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JAR-25 Large Aeroplanes

Section 2 - Acceptable Means of Compliance and Interpretations - ACJ

Change 14, 27 May 1994

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Section 2 - Acceptable Means of Compliance and Interpretations - ACJ

1 GENERAL
1.1 This Section contains Acceptable Means of Compliance and Interpretative Material that has been agreed for inclusion into JAR. These Acceptable Means of Compliance and Interpretations have been developed on the basis of the study of FAR 25 and their publication does not imply that the TSOs and Advisory Circulars associated with the Subparts of FAR 25 have been found acceptable as part of the JAR; consideration has still to be given to the acceptability of such documents.
2 PRESENTATION
2.1 The Acceptable Means of Compliance and Interpretative Material are presented in full page width on loose pages, each page being identified by the date of issue or the Change number under which it is amended or reissued.
2.2 A numbering system has been used in which the Acceptable Means of Compliance or Interpretative Material uses the same number as the paragraph in JAR to which it is related. The number is introduced by the letters ACJ (Advisory Circular - Joint) to distinguish the material from the JAR.
2.3 The nature of the advisory material is indicated immediately following the heading and for this purpose the two types of material are defined as follows:
Interpretative Material helps to illustrate the meaning of a requirement.
Acceptable Means of Compliance illustrate a means, but not the only means, by which a requirement can be met.
2.4 Explanatory Notes not forming part of the ACJ text appear in a smaller typeface.
[ 2.5 ] Subpart J - Gas Turbine Auxiliary Power Unit Installations - uses a similar numbering system to that used in Subpart E - Powerplant Installations, except that the number is prefaced either by the letter A or B, according to whether the requirement applies to all APUs or only to Essential APUs. This numbering system is continued in this Section 2, but to assist in the more rapid identification of ACJs relating to Subpart J, the words '(Auxiliary Power Units)' are added beside the title in each case.
[ 2.6 ] New, amended and corrected text is enclosed within heavy brackets.
ACJ - Subpart B

ACJ 25.21(d)
Proof of Compliance (Interpretative Material)
See JAR 25.21(d)
1 The tolerances specified in JAR 25.21(d) are the allowable deviations from the specified flight conditions for a particular test. They are not allowable tolerances on specific requirements, nor are they to be considered as allowable inaccuracies of measurement or of the method of determination. For example, when demonstrating static longitudinal stability with a specified trim speed of 14 VS1 the trim speed may be 14 VS1 3 knots or 3%; however, no positive tolerance is permitted when demonstrating the trim speed of not more than 14 VS1 required by JAR 25.161(c) (1).
2 Where variation of the parameter on which a tolerance is permitted will have an appreciable effect on the test, the result should be corrected to the specified value of the parameter; otherwise no correction is necessary.
3 In areas of critical handling or stability, notwithstanding the tolerance of JAR 25.21(d) (7% total travel), aft centre of gravity tests should be flown at a centre of gravity not more forward than the certificate aft centre of gravity limit. Tests which are critical on the forward centre of gravity limit should be flown at centres of gravity at least as forward as the certificate forward limit.

ACJ 25.101
General (Interpretative Material)
See JAR 25.101
The test aeroplane used in the determination of the scheduled performance should be in a condition which, as far as is reasonably possible, is representative of the average new production aeroplane. Where the test aeroplane differs from this standard (e.g. with regard to engine idle thrust settings, flap rigging, etc.) it will be necessary to correct the measured performance for any significant performance effects of such differences.

ACJ 25.101(c)
General (Interpretative Material)
[ See JAR 25.101(c) ]
1 In expanding flight test performance results to produce flight manual scheduled performance data, the following extrapolation limits and conservatism factors should normally be applied. These guide lines assume that the expansion will be made in the conventional way (i.e. using engine thrust data to derive an aeroplane drag polar and then to expand the results to cover the certification range). If any other expansion method is used the Authority should be consulted for appropriate extrapolation guide lines.
1.1 Take-off Performance
1.1.1 Weight. The variation of take-off performance with weight may be extrapolated without conservatism to a weight greater, by up to 10%, than the maximum weight tested and to a weight lower, by up to 10%, than the lowest weight tested. These ranges may not be applicable if there are significant discontinuities, or unusual variations, in the scheduling of take-off speeds with weight, in the weight ranges covered by extrapolation.
1.1.2 Altitude
a. Where the variation of engine thrust with altitude has been derived from acceptable altitude test facility testing, and has been confirmed by climb performance measurements over the required altitude range, the variation of take-off performance with altitude may be extrapolated by 6000 ft above the *W/d range tested, without conservatism. The take-off performance may be extrapolated beyond this 6000 ft figure provided a conservative margin of 2% is included in the scheduled distance, for each 1000 ft of extrapolation beyond 6000 ft.
b. In any other situation, altitude extrapolation without conservatism should be limited to 3000 ft and the above conservatism should be added for any further extrapolation.
*W = Weight
d = Ratio of ambient pressure to sea-level pressure.
1.1.3 Temperature
a. The variation of take-off performance with temperature may be extrapolated, without conservatism, to a temperature which exceeds the maximum temperature tested by -
i. 15 C or
ii. The amount by which the maximum temperature tested exceeds the minimum temperature tested, whichever is the greater.
b. The performance may also be extrapolated, without conservatism, by a similar amount below the minimum temperature tested.
c. The take-off performance may be extended beyond these limits provided a conservative margin of 2% is included in the scheduled distance for each 5C of extrapolation beyond the ranges quoted above.
d. The extent of permitted temperature extrapolation, or the magnitude of any margin of conservatism, where there are changes of engine limiting parameter (e.g. a change from pressure limit to TGT limit) in the range covered by the extrapolation, will depend on the availability of sufficient data from an appropriate test facility or heated inlet tests to allow extrapolation to be made with acceptable confidence.
1.2 Climb
1.2.1 Weight. As for take-off (see paragraph 1.1.1).
1.2.2 Altitude. Climb performance should normally be measured over a sufficiently large range of altitudes to warrant extrapolation without conservatism over the certificated range.
1.2.3 Temperature
a. Extrapolation ranges as for take-off (see paragraph 1.1.3).
b. For extrapolation beyond the ranges permitted without conservatism, the thrust assumed should be reduced by a conservative margin of 2% for each 5C of extrapolation beyond the 'without conservatism' limits.
1.3 Landing
1.3.1 Weight. As for take-off (see paragraph 1.1.1).

ACJ 25.101(h)(3)
General (Interpretative Material)
See JAR 25.101(h)(3)
There should be a minimum time delay of one second between operating each device which assists in the deceleration of the aeroplane motion in addition to, or in lieu of wheel brakes.
The minimum specified time delay should be applied as an additional margin to the actual time required for executing each procedure. However, if cockpit procedures require other crew members to operate devices on the captain's command, an additional time delay of one second should be required after the transmittal of each order to the other crew members by the captain.

ACJ 25.103(b)(1)
Stalling Speed (Interpretative Material)
[ See JAR 25.103(b)(1), and JAR 25.201(c)(1) ]
The stall entry rate is defined as the mean rate of speed reduction (in knots EAS/second) in the deceleration to the minimum speed achieved in the particular stall demonstration, from a speed 10% above that minimum speed, i.e.

Where Vmin is defined as the minimum speed achieved in the particular stall demonstration.

ACJ 25.103(c)
Stall Speed (Acceptable Means of Compliance)
See JAR 25.103(c)
In flight tests to determine stall speeds it is generally not possible to maintain a load factor of unity until the stall occurs. VS1g should be taken to be the minimum value of Vn obtained during the approach to the stall (up to the moment of stall identification), where V is the calibrated airspeed and n is the aeroplane load factor normal to the flight path, measured simultaneously.

ACJ 25.107(d)
Take-off Speeds (Acceptable Means of Compliance)
[ See JAR 25.107(d) ]
1 If cases are encountered where it is not possible to obtain the actual VMU at forward centre of gravity with aeroplanes having limited elevator power (including those aeroplanes which have limited elevator power only over a portion of the take-off weight range), it will be permissible to test with a more aft centre of gravity and/or more than normal nose-up trim to obtain VMU.
1.1 When VMU is obtained in this manner, the values should be corrected to those which would have been attained at forward centre of gravity if sufficient elevator power had been available. The variation of VMU with centre of gravity may be assumed to be the same as the variation of stalling speed in free air with centre of gravity for this correction.
1.2 In such cases where VMU has been measured with a more aft centre of gravity and/or with more than normal nose-up trim, the VR selected should (in addition to complying with the requirements of JAR 25.107(e)) be greater by an adequate margin than the lowest speed at which the nose wheel can be raised from the runway with centre of gravity at its most critical position and with the trim set to the normal take-off setting for the weight and centre of gravity.
NOTE: A margin of 5 knots between the lowest nose-wheel raising speed and VR would normally be considered to be adequate.
2 Take-offs made to demonstrate VMU should be continued until the aeroplane is out of ground effect. The aeroplane pitch attitude should not be decreased after lift-off.

ACJ 25.107(e)(1)(iv)
Take-off Speeds (Interpretative Material)
[ See JAR 25.107(e)(1)(iv) ]
1 In establishing the maximum practical rate of rotation the effect of centre of gravity position should be considered.
2 To be defined as a geometry limited aeroplane, the aeroplane should be geometry limited to the extent that a maximum gross weight take-off with the tail dragging will result [ in a clean lift-off and fly-away in the all-engines operating configuration. During such a take-off for all-engines-operating configurations, ] the resulting distance to the 35 ft height should not be greater than 105% of the normal take-off distance under similar weight, altitude and temperature conditions before the 15% margin is added. Lastly the VMU demonstrated should be sound and repeatable.
The criteria concerning the demonstration of the geometry limited proof test with regard to the capability for a clean lift-off and fly-away are as follows:
i. The aeroplane's pitch attitude from a speed 96% of the actual lift-off speed should be within 5% (in degrees) of the tail dragging attitude to the point of lift-off.
ii. During the above speed range (96% to 100% of the actual lift-off speed) the aft undersurface of the aeroplane should have achieved actual runway contact. It has been found practical in tests for contact to be maintained for at least 50% of the time that the aeroplane is in this speed range.
iii. It is acceptable for the undersurface of the test aeroplane to be modified in such a manner as to protect it from damage during these tests, provided that the effect on the limiting attitude (i.e. the pitch attitude at which the undersurface makes contact with the runway) is small.

ACJ 25.107(e)(3)
Take-off Speeds (Interpretative Material)
[ See JAR 25.107(e)(3) ]
In showing compliance with JAR 25.107(e)(3) -
a. Rotation at a speed of VR-5 knots should be carried out using, up to the point of lift-off, the same rotation technique, in terms of control input, as that used in establishing the one-engine-inoperative distance of JAR 25.113 (a)(1);
b. The engine failure speed used in the VR-5 knots demonstration should be the same as that used in the comparative take-off rotating at VR;
c. The tests should be carried out both at the lowest practical weight (such that VR-5 knots is not less than VMCG) and at a weight approaching take-off climb limiting conditions;
d. The tail or tail skid should not contact the runway.
ACJ No. 1 to JAR 25.107(e)(4)
Take-off Speeds (Interpretative Material)
[ See JAR 25.107(e)(4) ]
Reasonably expected variations in service from established take-off procedures should be evaluated in respect of out-of-trim conditions during certification flight test programmes. For example, normal take-off should be made with the longitudinal control trimmed to its most adverse position within the allowable take-off trim band.
ACJ No. 2 to JAR 25.107(e)(4)
Take-off Speeds (Interpretative Material)
[ See JAR 25.107(e)(4) ]
[ 1 JAR 25.107(e)(4) states that there must be no marked increase in the scheduled take-off distance ] when reasonably expected service variations, such as over-rotation, are encountered. This can be interpreted as requiring take-off tests with all engines operating with an abuse on rotation speed.
2 The expression 'marked increase' in the take-off distance is defined as any amount in excess of 1% of the scheduled take-off distance. Thus the abuse test should not result in a field length more than 101% of the scheduled field length.
3 For the early rotation abuse condition with all engines operating and at a weight as near as practicable to the maximum sea-level take-off weight, it should be shown by test that when the aeroplane is rotated rapidly at a speed which is 7% or 10 knots, whichever is lesser, below the scheduled VR speed, no 'marked increase' in the scheduled field length would result.

ACJ 25.109(a)
Accelerate-stop Distance (Interpretative Material)
[ See JAR 25.109(a) ]
Propeller pitch position. When conducting accelerate-stop tests on propeller-driven aeroplanes, it is permissible to place the propellers of the operating engines in the ground idle pitch setting, with a suitable time delay after engine failure (see also ACJ 25.101(h)(3)) and provided that the aeroplane remains controllable on a slippery runway. The propeller on the inoperative engine should be in the position it would normally take up upon closing all power levers.

ACJ 25.111
Take-off Path (Interpretative Material)
[ See JAR 25.111 ]
[ The height references in JAR 25.111 should be interpreted as geometrical heights. ]

ACJ 25.111(b)
Take-off Path (Interpretative Material)
[ See JAR 25.111(b) ]
1 Rotation speed, VR, is intended to be the speed at which the pilot initiates action to raise the nose gear off the ground, during the acceleration to V2; consequently, the take-off path determination, in accordance with JAR 25.111 (a) and (b), should assume that pilot action to raise the nose gear off the ground will not be initiated until the speed VR has been reached.
2 The time between lift-off and the initiation of gear retraction should be not less than 3 seconds and may need to be longer than 3 seconds if, on a particular aeroplane type, a longer delay is found to be appropriate.

ACJ 25.113(a)(2)
Take-off Distance and Take-off Run (Interpretative Material)
[ See JAR 25.113(a)(2), JAR 25.113(b)(2) ]
In establishment of the take-off distance and take-off run, with all engines operating, in accordance with JAR 25.113(a) and (b), the flight technique should be such that -
a. A speed of not less than V2 is achieved before reaching a height of 35 ft above the take-off surface,
b. It is consistent with the achievement of a smooth transition to a steady initial climb speed of not less than V2 + 10 knots at a height of 400 ft above the take-off surface.

ACJ 25.119(a)
Landing Climb: All-engines-operating (Interpretative Material)
[ See JAR 25.119(a) ]
[ In establishing the thrust specified in JAR 25.119(a), either - ]
a. Engine acceleration tests should be conducted using the most critical combination of the following parameters:
i. Altitude;
ii. Airspeed;
iii. Engine bleed;
iv. Engine power off-take;
likely to be encountered during an approach to a landing airfield within the altitude range for which landing certification is sought; or
[ b. The thrust specified in JAR 25.119(a) should be established as a function of these parameters. ]

ACJ 25.121
Climb: One-engine-inoperative (Acceptable Means of Compliance and Interpretative Material)
[ See JAR 25.121 ]
[ 1 In showing compliance with JAR 25.121 it is accepted that bank angles of up to 2 to 3 toward the ] operating engine(s) may be used.
[ 2 The height references in JAR 25.121 should be interpreted as geometrical heights. ]
[ ACJ 25.121(a)
Climb: One-engine-inoperative (Interpretative Material)
See JAR 25.121(a)
The configuration of the landing gear used in showing compliance with the climb requirements of JAR 25.121(a) may be that finally achieved following "gear down" selection. ]

ACJ 25.121(a)(1)
Climb: One-engine-inoperative (Interpretative Material)
See JAR 25.121(a)(1)
A 'power operating condition' more critical than that existing at the time when retraction of the landing gear is begun would occur, for example, if water injection were discontinued prior to reaching the point at which the landing gear is fully retracted.

ACJ 25.121(b)(1)
Climb: One-engine-inoperative (Interpretative Material)
See JAR 25.121(b)(1)
A 'power operating condition' more critical than that existing at the time the landing gear is fully retracted would occur, for example, if water injection were discontinued prior to reaching a gross height of 400 ft.

ACJ 25.123
En-route Flight Paths (Interpretative Material)
See JAR 25.123
If, in showing compliance with JAR 25.123, any credit is to be taken for the progressive use of fuel by the operating engines, the fuel flow rate should be assumed to be 80% of the engine specification flow rate at maximum continuous power, unless a more appropriate figure has been substantiated by flight tests.
[ ACJ 25.125(a)(2)
Landing Approach Speed (Acceptable Means of Compliance)
See JAR 25.125(a)(2)
See Orange Paper Amendment 96/1
If VMCL exceeds 13 VS at any practicable landing weight, the landing distance should be determined for an approach speed not less than the greater of VMCL and 13 VS. Practicable landing weight is defined as empty weight plus fuel for 100 n.m. diversion and 30 minute hold, minimum flight and cabin crew and 10% of maximum payload. ]

ACJ 25.125(a)(3)
Change of Configuration (Interpretative Material)
See JAR 25.125(a)(3)
No changes in configuration, addition of thrust, or nose depression should be made after reaching 50 ft height.
[ ACJ 25.125(b) ]
Landing (Interpretative Material)
[ See JAR 25.125(b) ]
1 During measured landings, if the brakes can be consistently applied in a manner permitting the nose gear to touch down safely, the brakes may be applied with only the main wheels firmly on the ground. Otherwise, the brakes should not be applied until all wheels are firmly on the ground.
2 This is not intended to prevent operation in the normal way of automatic braking systems which, for instance, permit brakes to be selected on before touchdown.
[ ACJ 25.125(b)(2) ]
Landing (Interpretative Material)
[ See JAR 25.125(b)(2) ]
To ensure compliance with JAR 25.125(b)(2), a series of six measured landings should be conducted on the same set of wheel brakes and tyres.

ACJ 25.143(a) and (b)
Controllability and Manoeuvrability (Interpretative Material)
See JAR 25.143(a) and (b)
In showing compliance with the requirements of JAR 25.143(a) and (b) account should be taken of aeroelastic effects and structural dynamics (including aeroplane response to rough runways and water waves) which may influence the aeroplane handling qualities in flight and on the surface. The oscillation characteristics of the flightdeck, in likely atmospheric conditions, should be such that there is no reduction in ability to control and manoeuvre the aeroplane safely.

ACJ 25.143(b)(1)
Control Following Engine Failure (Acceptable Means of Compliance)
See JAR 25.143(b)(1)
1 An acceptable means of showing compliance with JAR 25.143(b)(1) is to demonstrate that it is possible to regain full control of the aeroplane without attaining a dangerous flight condition in the event of a sudden and complete failure of the critical engine in the following conditions:
a. At each take-off flap setting at the lowest speed recommended for initial steady climb with all engines operating after take-off, with -
i. All engines, prior to the critical engine becoming inoperative, at maximum take-off power or thrust;
ii. All propeller controls in the take-off position;
iii. The landing gear retracted;
iv. The aeroplane in trim in the prescribed initial conditions; and
b. With wing-flaps retracted at a speed of 13 VS1 with -
i. All engines, prior to the critical engine becoming inoperative, at maximum continuous power or thrust;
ii. All propeller controls in the en-route position;
iii. The landing gear retracted;
iv. The aeroplane in trim in the prescribed initial conditions.
2 The demonstrations should be made with simulated engine failure occurring during straight flight with wings level. In order to allow for likely delay in the initiation of recovery action, no action to recover the aeroplane should be taken for 2 seconds following engine failure. The recovery action should not necessitate movement of the engine, propeller or trimming controls, nor require excessive control forces. The aeroplane will be considered to have reached an unacceptable attitude if a bank angle of 45 is exceeded during recovery.

ACJ 25.143 (c)
Controllability and Manoeuvrability (Interpretative Material)
See JAR 25.143 (c)
See Orange Paper Amendment 96/1
1 The maximum forces given in the table in JAR 25.143(c) for pitch and roll control for temporary application are applicable to manoeuvres in which the control force is only needed for a short period. Where the manoeuvre is such that the pilot will need to use one hand to operate other controls (such as the landing flare, or changes of configuration or power which result in a change of control force which must be trimmed out) the single-handed maximum control forces will be applicable. In other cases (such as take-off rotation, or manoeuvring during en-route flight) the two-handed maximum forces will apply.
2 The maximum forces for prolonged application are intended to apply to periods in excess of approximately 10 minutes. For periods shorter than 10 minutes, reasonable interpolation between the 'temporary' and 'prolonged' forces will be permitted.
3 These maximum forces apply under probable operating conditions (see JAR 25.143(b)). In the event [ of failure conditions which are assessed as Improbable (see AMJ 25.1309) greater forces may be acceptable. ]
ACJ No. 1 to JAR 25.143 (f)
Controllability and Manoeuvrability (Acceptable Means of Compliance)
See JAR 25.143 (f)
See Orange Paper Amendment 96/1
An acceptable means of compliance with the requirement that stick forces may not be excessive when manoeuvring the aeroplane, is to demonstrate that, in a turn for 05g incremental normal acceleration (03g above 20 000 ft (6096 m)) at speeds up to VMO/MMO, the temporary two-handed control force of JAR 25.143(c) is not exceeded.
ACJ No. 2 to JAR 25.143(f)
Controllability and Manoeuvrability (Interpretative Material)
See JAR 25.143(f)
See Orange Paper Amendment 96/1
1 The aircraft will be considered to have been overstressed if limit strength has been exceeded in any critical component. For the purpose of this ACJ, limit strength is defined as the minimum demonstrated strength against the relevant manoeuvre load condition divided by 15.
2 Minimum Stick Force to Reach Limit Strength
2.1 The stick force necessary to reach limit strength in steady manoeuvre or wind up turns should not be less than 50 pounds, except that if severe buffeting occurs before the limit strength condition is reached a somewhat lower stick force may be acceptable. This minimum stick force applies in the en-route configuration with the aeroplane trimmed for straight flight, at all speeds from the minimum speed at which the limit strength condition can be achieved without stalling. No minimum stick force is specified for other configurations, but the requirements of JAR 25.143(f) are applicable in these conditions.
2.2 The acceptability of a stick force of less than 50 pounds at the limit strength condition will depend upon the intensity of the buffet, the adequacy of the warning margin (i.e. the load factor increment between buffet onset and the limit strength condition) and also on the acceptability of any associated non-linearities of the stick force characteristics.
3 Stick Force Characteristics
3.1 At all points within the buffet onset boundary determined in accordance with JAR 25.251 (e), but not including speeds above VMO/MMO (see JAR 25.253(a)(3)), the stick force should increase progressively with increasing load factor. Any reduction in stick force gradient with change of load factor should not be so large or abrupt as to impair significantly the ability of the pilot to maintain precise control over the load factor and pitch attitude of the aeroplane.
3.2 Beyond the buffet onset boundary hazardous stick force characteristics should not be encountered within the permitted manoeuvring envelope without adequate prior warning being given by severe buffeting or high stick forces. It should at all times be possible, by use of the primary longitudinal control alone, to pitch the aeroplane rapidly nose down so as to regain the initial trimmed conditions. The stick force characteristics demonstrated should comply with the following:
a. For normal acceleration increments of up to 03 g beyond buffet onset, where these can be achieved, the ability to control pitch attitude and load factor with precision should be retained. Local reversal of the stick force gradient within this range of load factor will be acceptable provided that any tendency to pitch up is mild and easily controllable.
b. For normal acceleration increments of more than 03 g beyond buffet onset, where these can be achieved, more marked reversals of the stick force gradient may be acceptable. It should be possible for any tendency to pitch up to be contained within the allowable manoeuvring limits without applying push forces to the control column and without making large and rapid forward movement of the control column.
[ 3.3 In flight tests to satisfy paragraph 3.1 and 3.2 the load factor should be increased until either - ]
a. The level of buffet becomes sufficient to provide an obvious warning to the pilot which is a strong deterrent to further application of load factor; or
b. Further increase of load factor requires a stick force in excess of 150 pounds (or in excess of 100 pounds when beyond the buffet onset boundary) or is impossible because of the limitations of the control system; or
c. The positive limit manoeuvring load factor established in compliance with JAR 25.337(b) is achieved.
4 Negative Load Factors
It is not intended that a detailed flight test assessment of the manoeuvring characteristics under negative load factors should necessarily be made throughout the specified range of conditions. An assessment of the characteristics in the normal flight envelope involving normal accelerations from 1 g to 0 g, will normally be sufficient. Stick forces should also be assessed during other required flight testing involving negative load factors. Where these assessments reveal stick force gradients that are unusually low, or that are subject to significant variation, a more detailed assessment, in the most critical of the specified conditions, will be required. This may be based on calculations provided these are supported by adequate flight test or wind tunnel data.

ACJ 25.145 (b)(2)
Longitudinal Control (Interpretative Material)
See JAR 25.145 (b)(2)
Where high lift devices are being retracted and where large and rapid changes in maximum lift occur as a result of movement of high-lift devices, some reduction in the margin above the stall may be accepted.

ACJ 25.145(d)
Longitudinal Control - Take-off Climb (Acceptable Means of Compliance)
See JAR 25.145 (d)
An acceptable method of demonstrating compliance with the requirement of JAR 25.145(d) would be to demonstrate that the speeds used to show compliance with JAR 25.121(b) and (c) and JAR 25.111(c)(3) are not less than 108 times the respective speeds at which stall warning first occurs under the same conditions of configuration and power. This may be demonstrated in still air and with a slow reduction of speed until stall warning is encountered. Where natural pre-stall buffeting is followed at a lower speed by artificial stall warning which satisfies the requirements of JAR 25.207, the natural buffeting should, nevertheless be considered to constitute stall warning for the purposes of this requirement if it is of such an intensity as to make it likely that a pilot would take action to increase speed.
[ ACJ 25.147 (a)
Directional Control; general (Interpretative Material)
See JAR 25.147 (a)
The intention of the requirement is that the aircraft can be yawed as prescribed without the need for application of bank angle. Small variations of bank angle that are inevitable in a realistic flight test demonstration are acceptable. ]

ACJ 25.147 (c)(2)
Lateral Control: One Engine Inoperative (Interpretative Material)
See JAR 25.147(c)(2)
An acceptable method of demonstrating compliance with JAR 25.147(c)(2) is as follows:
It should be possible in the conditions specified below to roll the aeroplane from a steady 30 banked turn through an angle of 60 so as to reverse the direction of the turn in not more than 11 seconds. In this demonstration the rudder may be used to the extent necessary to minimise sideslip. The demonstration should be made rolling the aeroplane in either direction, and the manoeuvre may be unchecked.
Conditions: Airspeed V2.
Wing-flaps. In each take-off position.
Landing Gear. Retracted.
Power. The critical engine inoperative and its propeller (if applicable) in the minimum drag condition; the remaining engines operating at maximum take-off power.
Trim. The aeroplane should be in trim, or as nearly as possible in trim, for straight flight in these conditions, and the trimming controls should not be moved during the manoeuvre.

ACJ 25.147(e)
Lateral Control: All Engines Operating (Interpretative Material)
See JAR 25.147(e)
An acceptable method of demonstrating that roll response and peak roll rates are adequate for compliance with JAR 25.147 (e) is as follows:
It should be possible in the conditions specified below to roll the aeroplane from a steady 30 banked turn through an angle of 60 so as to reverse the direction of the turn in not more than 7 seconds. In these demonstrations the rudder may be used to the extent necessary to minimise sideslip. The demonstrations should be made rolling the aeroplane in either direction, and the manoeuvres may be unchecked.
Conditions:
(a) En-route: Airspeed. All speeds between the minimum value of the scheduled all-engines-operating climb speed and VMO/MMO .
Wing-flaps. En-route position(s).
Air Brakes. All permitted settings from Retracted to Extended.
Landing Gear. Retracted.
Power. All engines operating at all powers from flight idle up to maximum continuous power.
Trim. The aeroplane should be in trim from straight flight in these conditions, and the trimming controls should not be moved during the manoeuvre.
(b) Approach: Airspeed. Either the speed maintained down to the 50 ft height in compliance with JAR 25.125(a)(2), or the target threshold speed determined in accordance with JAR 25.125 (c)(2)(i) as appropriate to the method of landing distance determination used.
Wing-flaps. In each landing position.
Air Brakes. In the maximum permitted extended setting.
Landing Gear. Extended.
Power. All engines operating at the power required to give a gradient of descent of 50%.
Trim. The aeroplane should be in trim for straight flight in these conditions, and the trimming controls should not be moved during the manoeuvre.

ACJ 25.149
Minimum Control Speeds (Interpretative Material)
[ See JAR 25.149 ]
1 The determination of the minimum control speed, VMC, and the variation of VMC with available thrust, may be made primarily by means of 'static' testing, in which the speed of the aeroplane is slowly reduced, with the thrust asymmetry already established, until the speed is reached at which straight flight can no longer be maintained. A small number of 'dynamic' tests, in which sudden failure of the critical engine is simulated, should be made in order to check that the VMCs determined by the static method are valid.
2 When minimum control speed data are expanded for the determination of minimum control speeds (including VMC, VMCG AND VMCL) for all ambient conditions, these speeds should be based on the maximum values of thrust which can reasonably be expected from a production engine in service. The minimum control speeds should not be based on specification thrust, since this thrust represents the minimum thrust as guaranteed by the manufacturer, and the resulting speeds would be unconservative for most cases.

ACJ 25.149(e)
Minimum Control Speed (Interpretative Material)
[ See JAR 25.149(e) ]
See Orange Paper Amendment 96/1
When carrying out tests for the determination of VMCG the pilot conducting the tests should recognise engine failure by external reference only and not by reference to the engine instruments. The pilot conducting the tests should not be aware of which engine is to be made inoperative, nor of the precise speed at which this is to occur.

ACJ 25.149(f)
See Orange Paper Amendment 96/1

ACJ 25.149(g)
See Orange Paper Amendment 96/1

ACJ 25.149(h)(3)
See Orange Paper Amendment 96/1

ACJ 25.149(h)(4)
See Orange Paper Amendment 96/1

ACJ 25.173(c)
Static Longitudinal Stability (Interpretative Material)
See JAR 25.173(c)
The average gradient is taken over each half of the speed range between 085 and 115 Vtrim.

ACJ 25.177(b)
Static Lateral Stability (Interpretative Material)
[ See JAR 25.177(b) ]
1 For speeds between 12 VS1 and 13 VS1 for wing-flap positions more extended than the most extended take-off wing-flap setting, the symmetric power used during demonstrations need not exceed the power required for level flight in the conditions (speed and configuration) in which the demonstration is made.
[2 Demonsrtation of compliance with Jar 25.177(b) should be made from sideslip angles appropriate] to the operation of the aeroplane. Sideslip angles corresponding to half rudder deflection would normally be considered appropriate for this purpose.
3 The requirement is concerned with the short-term response of the aeroplane, and long term effects, due to factors such as fuel movement, need not be taken into account. If the initial response of the aeroplane on releasing the aileron control is neutral this will be acceptable, even through the response gradually becomes unstable in the longer term.

ACJ 25.177(c)
Static Directional and Lateral Stability (Interpretative Material)
See JAR 25.177(c)
For the range of sideslip angles appropriate to the operation of the aeroplane the speed range covered should be as specified in JAR 25.177(a). For greater angles the tests should cover all permitted or scheduled operational speeds, configurations and associated power settings.

ACJ 25.181
Dynamic Stability (Interpretative Material)
[ See JAR 25.181 ]
The requirements of JAR 25.181 are applicable at all speeds between the stalling speed and VFE, VLE or VFC/MFC, as appropriate.

ACJ 25.201(a)(2)
Stall Demonstration (Interpretative Material)
See JAR 25.201(a)(2)
The power for all power-on stall demonstrations is that power necessary to maintain level flight at a speed of 16 VS1 at maximum landing weight, with flaps in the approach position and landing gear retracted, where VS1 is the stall speed in the same conditions (except with idle power). The flap position to be used to determine this power setting is that position in which the stall speed does not exceed 110% of the stall speed with the flaps in the most extended landing position.

ACJ 25.201(b)(1)
Stall Demonstration (Interpretative Material)
See JAR 25.201(b)(1)
Stall demonstrations for compliance with JAR 25.201 should include demonstrations with deceleration devices deployed for all flap positions unless limitations against use of the devices with particular flap positions are imposed. 'Deceleration devices' include spoilers when used as air brakes, and thrust reversers when use in flight is permitted. Stall demonstrations with deceleration devices deployed should normally be carried out with power off, except where deployment of the deceleration devices while power is applied is likely to occur in normal operations (e.g. use of extended air brakes during landing approach).

ACJ 25.201(c)(2)
See Orange Paper Amendment 96/1

ACJ 25.201(c)(3)
Stall Demonstration (Acceptable Means of Compliance)
See JAR 25.201(c)(3)
See Orange Paper Amendment 96/1
An acceptable test procedure for dynamic stalls is as follows:
a. Establish the aeroplane in a 30 banked turn at a speed between 13 VS and 14 VS.
b. Commence reducing speed at approximately 1 knot per second.
c. At a speed between 13 VS and the speed at which stall warning commences, increase the rate of change of angle of attack so that the stall occurs at a normal acceleration of between 13 g and 15 g. Maintain the increased rate of change of angle of attack until the stall occurs.
d. When the aeroplane has stalled, recover by normal recovery techniques.

ACJ 25.201(d)
Stall Demonstration (Interpretative Material)
See JAR 25.201(d)
1 The behaviour of the aeroplane includes the behaviour as affected by the normal functioning of any systems with which the aeroplane is equipped, including devices intended to alter the stalling characteristics of the aeroplane.
2 Unless the design of the automatic flight control system of the aeroplane protects against such an event, the stalling characteristics and adequacy of stall warning, when the aeroplane is stalled under the control of the automatic flight control system, should be investigated. (See also JAR 25.1329(f).)

ACJ 25.201(d)(3)
See Orange Paper Amendment 96/1

ACJ 25.203
Stall Characteristics (Interpretative Material)
See JAR 25.203
1 Static Longitudinal Stability during the Approach to the Stall. During the approach to the stall the longitudinal control pull force should increase continuously as speed is reduced from the trimmed speed to approximately 11 VS. At speeds below 11 VS some reduction in longitudinal control pull force will be acceptable provided that it is not sudden or excessive.
2 Rolling Motions at the Stall
2.1 In level wing stalls the bank angle may exceed 20 occasionally, provided that lateral control is effective during recovery.
2.2 For stalls from a 30 banked turn with an entry rate of 1 knot per second, the maximum bank angle which occurs during the recovery should not exceed approximately 60 in the original direction of the turn, or 30 in the opposite direction.
2.3 For dynamic stalls the maximum bank angle which occurs during the recovery should not exceed approximately 90 in the original direction of the turn, or 60 in the opposite direction.
3 Deep Stall Penetration. Where the results of wind tunnel tests reveal a risk of a catastrophic phenomenon (e.g. superstall, a condition at angles beyond the stalling incidence from which it proves difficult or impossible to recover the aeroplane), studies should be made to show that adequate recovery control is available at and sufficiently beyond the stalling incidence to avoid such a phenomenon.

ACJ 25.207(b)
Stall Warning (Interpretative Material)
See JAR 25.207(b)
1 A warning which is clear and distinctive to the pilot is one which cannot be misinterpreted or mistaken for any other warning, and which, without being unduly alarming, impresses itself upon the pilot and captures his attention regardless of what other tasks and activities are occupying his attention and commanding his concentration. Where stall warning is to be provided by artificial means, a stick shaker device producing both a tactile and an audible warning is an Acceptable Means of Compliance.
2 Where stall warning is provided by means of a device, compliance with the requirement of JAR 25.21(e) should be established by ensuring that the device has a high degree of reliability. One means of complying with this criterion is to provide dual independent systems.

ACJ 25.251(e)
Vibration and Buffeting in Cruising Flight (Acceptable Means of Compliance)
See JAR 25.251(e)
1 Probable Inadvertent Excursions beyond the Buffet Boundary
1.1 JAR 25.251(e) states that probable inadvertent excursions beyond the buffet onset boundary may not result in unsafe conditions.
1.2 An acceptable means of compliance with this requirement is to demonstrate by means of flight tests beyond the buffet onset boundary that hazardous conditions will not be encountered within the permitted manoeuvring envelope (as defined by JAR 25.337) without adequate prior warning being given by severe buffeting or high stick forces.
1.3 Buffet onset is the lowest level of buffet intensity consistently apparent to the flight crew during normal acceleration demonstrations in smooth air conditions.
1.4 In flight tests beyond the buffet onset boundary to satisfy paragraph 1.2, the load factor should be increased until either -
a. The level of buffet becomes sufficient to provide an obvious warning to the pilot which is a strong deterrent to further application of load factor; or
b. Further increase of load factor requires a stick force in excess of 100 pounds, or is impossible because of the limitations of the control system; or
c. The positive limit manoeuvring load factor established in compliance with JAR 25.337(b) is achieved.
1.5 Within the range of load factors defined in paragraph 1.4 no hazardous conditions (such as hazardous involuntary changes of pitch or roll attitude, engine or systems malfunctioning which require urgent corrective action by the flight crew, or difficulty in reading the instruments or controlling the aeroplane) should be encountered.
2 Range of Load Factor for Normal Operations
2.1 JAR 25.251(e) requires that the envelopes of load factor, speed, altitude and weight must provide a sufficient range of speeds and load factors for normal operations.
2.2 An acceptable means of compliance with the requirement is to establish the maximum altitude at which it is possible to achieve a positive normal acceleration increment of 03 g without exceeding the buffet onset boundary. See also ACJ 25.1585(c).

ACJ 25.253(a)(3)
Lateral Control at Speeds in Excess of VMO/MMO (Interpretative Material)
See JAR 25.253(a)(3)
An acceptable method of demonstrating that roll capability is adequate to assure prompt recovery from a laterally upset condition is as follows:
It should be possible using lateral control alone to roll the aeroplane from a steady 20 banked turn through an angle of 40 so as to reverse the direction of the turn in not more than 8 seconds. The demonstration should be made rolling the aeroplane in either direction in the conditions specified below. The manoeuvres may be unchecked.
Conditions: Air Speed. All speeds from VMO/MMO to a speed close to VDF/MDF limited to the extent necessary to accomplish the manoeuvre and recovery without exceeding VDF/MDF.
Wing-flaps. En-route position(s).
Air Brakes. All permitted settings from retracted to extended.
Landing Gear. Retracted.
Power (i) All engines operating at the power required to maintain level flight at VMO/MMO, except that maximum continuous power need not be exceeded; and
(ii) If the effect of power is significant, with the throttles closed.
Trim. The aeroplane trimmed for straight flight at VMO/MMO. The trimming controls should not be moved during the manoeuvre.
In addition it should be demonstrated that use of rudder in the conventional sense, if it results in an adverse effect on roll rate, will not result in a dangerous reduction in lateral control capability.

ACJ 25.253(a)(5)
High Speed Characteristics (Interpretative Material)
See JAR 25.253(a)(5)
1 Trim Change due to Airbrake Selection
1.1 A load factor should be regarded as excessive if it exceeds 20 with stick free.
1.2 A nose-down pitching moment should be regarded as small if it necessitates a stick-force of less than 20 lb to maintain 1 g flight.
1.3 Compliance with the requirements of JAR 25.253(a)(5) should be demonstrated at speeds up to a speed close to VDF/MDF but limited to the extent necessary to accomplish the manoeuvre and recovery without exceeding VDF/MDF.

ACJ 25.255
Out-of-trim Characteristics (Interpretative Material)
[ See JAR 25.255 ]
1 Amount of Out-of-trim Required
1.1 The equivalent degree of trim, specified in JAR 25.255(a)(1) for aeroplanes which do not have a power-operated longitudinal trim system, has not been specified in quantitative terms, and the particular characteristics of each type of aeroplane must be considered. The intent of the requirement is that a reasonable amount of out-of-trim should be investigated, such as might occasionally be applied by a pilot.
1.2 In establishing the maximum mistrim that can be sustained by the autopilot the normal operation of the autopilot and associated systems should be taken into consideration. Where the autopilot is equipped with an auto-trim function the amount of mistrim which can be sustained will generally be small or zero. If there is no auto-trim function, consideration should be given to the maximum amount of out-of-trim which can be sustained by the elevator servo without causing autopilot disconnect.
2 Datum Trim Setting
2.1 For showing compliance with JAR 25.255(b)(1) for speeds up to VMO/MMO, the datum trim setting should be the trim setting required for trimmed flight at the particular speed at which the demonstration is to be made.
2.2 For showing compliance with JAR 25.255(b)(1) for speeds from VMO/MMO to VFC/MFC, and for showing compliance with JAR 25.255(b)(2) and (f), the datum trim setting should be the trim setting required for trimmed flight at VMO/MMO.
3 Reversal of Primary Longitudinal Control Force at Speeds greater than VFC/MFC
3.1 JAR 25.255(b)(2) requires that the direction of the primary longitudinal control force may not reverse when the normal acceleration is varied, for +1 g to the positive and negative values specified, at speeds above VFC/MFC. The intent of the requirement is that it is permissible that there is a value of g for which the stick force is zero, provided that the stick force versus g curve has a positive slope at that point (see Figure 1).


FIGURE 1
[ 3.2 If stick force characteristics are marginally acceptable, it is desirable that there should be no reversal of normal control sensing, i.e. an aft movement of the control column should produce an aircraft motion in the nose-up direction and a change in aircraft load factor in the positive direction, and a forward movement of the control column should change the aircraft load factor in the negative direction.
3.3 It is further intended that reversals of direction of stick force with negative stick-force gradients should not be permitted in any mistrim condition within the specified range of mistrim. If test results indicate that the curves of stick force versus normal acceleration with the maximum required mistrim have a negative gradient of speeds above VFC/MFC then additional tests may be necessary. The additional tests should verify that the curves of stick force versus load factor with mistrim less than the maximum required do not unacceptably reverse, as illustrated in the upper curve of Figure 2. Control force characteristics as shown in Figure 3, may be considered acceptable, provided that the control sensing does not reverse (see paragraph 3.2).

FIGURE 2 FIGURE 3
4 Probable Inadvertent Excursions beyond the Boundaries of the Buffet Onset Envelopes. JAR 25.255(e) states that manoeuvring load factors associated with probable inadvertent excursions beyond the boundaries of the buffet onset envelopes determined under JAR 25.251(e) need not be exceeded. It is intended that test flights need not be continued beyond a level of buffet which is sufficiently severe that a pilot would be reluctant to apply any further increase in load factor.
5 Use of the Longitudinal Trim System to Assist Recovery
5.1 JAR 25.255(f) requires the ability to produce at least 1.5 g for recovery from an overspeed condition of VDF/MDF, using either the primary longitudinal control alone or the primary longitudinal control and the longitudinal trim system. Although the longitudinal trim system may be used to assist in producing the required normal acceleration, it is not acceptable for recovery to be completely dependent upon the use of this system. It should be possible to produce 12 g by applying not more than 125 pounds of longitudinal control force using the primary longitudinal control alone.
5.2 Recovery capability is generally critical at altitudes where airspeed (VDF) is limiting. If at higher altitudes (on the MDF boundary) the manoeuvre capability is limited by buffeting of such an intensity that it is a strong deterrent to further increase in normal acceleration, some reduction of manoeuvre capability will be acceptable, provided that it does not reduce to below 13 g. The energy speed for flight test demonstrations of compliance with this requirement should be limited to the extent necessary to accomplish a recovery without exceeding VDF/MDF, and the normal acceleration should be measured as near to VDF/MDF as is practical. ]

ACJ 25.261
Flight in Rough Air (Interpretative Material)
See JAR 25X261
The procedures should give the maximum protection against loss of control and against structural damage occurring either directly as the result of turbulence or occurring in the recovery from any disturbance of the flight path. The procedures should, where necessary, distinguish also between the procedure to be followed when deliberately entering an area of known turbulence and that to be followed when the encounter is unexpected.
ACJ - Subpart C

ACJ 25.301(b)
Loads (Interpretative Material)
See JAR 25.301(b) and JAR 25.361(d)
The engine and its mounting structure are to be stressed to the loading cases for the aeroplane as a whole, including manoeuvring and gust loading conditions, together with conservative estimates of torque thrust, gyroscopic loading and any loading which may result from engine fans. Full allowance should be made for structural flexibility effects in landing cases. This also applies to auxiliary power units.

ACJ 25.307
Proof of Structure (Interpretative Material)
See JAR 25.307
In deciding the need for and the extent of testing including the load levels to be achieved the following factors will be considered by the Authority.
a. The confidence which can be attached to the constructors' overall experience in respect to certain types of aeroplanes in designing, building and testing aeroplanes.
b. Whether the aeroplane in question is a new type or a development of an existing type having the same basic structural design and having been previously tested, and how far static strength testing can be extrapolated to allow for development of the particular type of aeroplane.
c. The importance and value of detail and/or component testing including representation of parts of structure not being tested, and
d. The degree to which credit can be given for operating experience where it is a matter of importing for the first time an old type of aeroplane which has not been tested.

ACJ 25.331(c)(2)
Checked Manoeuvre (Acceptable Means of Compliance)
See JAR 25.331(c)(2)
The aeroplane being initially in balanced flight at n=1 at any speed between VA and VD, checked pitch manoeuvres should be studied up to the proof factors (n1 and Og) the load factors being maximum values obtained under transient conditions. It will be assumed that the manoeuvres meet the following description:
The elevator is moved rapidly in one direction then in the other to a position well beyond the original position before returning to that position: h = ho sin wt, h being the control surface deflection angle and w being the circular frequency of the angular movement of the control surface, taken to be equal to the undamped natural frequency of the short period rigid mode, but not being less than -

VA is the manoeuvring speed, and V the speed in question.
In general, it will only be necessary to analyse three quarters of the movement, assuming that the return of the control is effected in a less sudden manner.
The speed of movement of the pitching control specified above whilst maintaining the maximum normal acceleration to be achieved in the manoeuvre, may be adjusted to take into account limitations which may be imposed by maximum pilot effort specified in paragraph JAR 25.397(b), control system stops and any indirect effect imposed by limitations in the output side of the control system, such as stalling torque or maximum rate obtainable by a power control system.
[ ACJ 25.335(b)(2)
Design Diving Speed (Interpretative Material)
See JAR 25.335(b)(2)
In the absence of evidence supporting alternative criteria, compliance with JAR 25.335(b)(2) can be shown by provision of a margin between VC/MC and VD/MD sufficient to provide for the following atmospheric conditions:
1. Encounter with a Horizontal Gust. The effect of encounters with a substantially head-on draught, which is assumed to act at the most adverse angle between 30 above and 30 below the flight path, will be considered. The draught velocity will be 50 ft/s EAS at altitudes up to 20 000 ft; at altitudes above 20 000 ft the draught velocity may be reduced linearly from 50 ft/s at 20 000 ft to 25 ft/s at 50 000 ft, above which altitude the draught velocity will be considered constant. The draught will be assumed to build up in not more than 2 seconds and to last for 30 seconds.
2. Entry into Jet Streams or Regions of High Wind Shear
2.1 Horizontal and Vertical wind shears will be investigated taking into account the wind shear data of this paragraph which are world-wide extreme values.
NOTE: The high inertia of any fast moving vehicle means that the airspeed will change if the vehicle enters a region of wind shear. The extent of this change is a function of the rate at which the vehicle enters the wind shear, the mode of control of the vehicle and its aerodynamic characteristics. In certain cases, the control characteristics may result in over corrections resulting in a speed change once the vehicle leaves the region of wind shear.
2.2 With the aeroplane being operated in a normal manner (i.e. at normal rates of climb and descent), the worst wind shears which it might encounter, according to available meteorological data, can be expressed as follows:
a. Horizontal Wind Shear The jet stream is assumed to have the following characteristics:
A linear shear on one side of the core of 0001 sec -1 (36 kt/NM) over a distance of 25 NM or of 00007 sec -1 (252 kt/NM) over a distance of 50 NM or of 00005 sec -1 (18 kt/NM) over a distance of 100 NM whichever is the more severe.
Both cases of positive and negative shear should be considered, and the shear on the other side of the core may be taken as 2/3 of the primary value, but of opposite sign.
b. Vertical Wind Shear The jet stream is assumed to have the following characteristics:
The shear above the core is taken as equal to the shear below the core, but of opposite sign, and is the most severe of the cases given in Table 1. Cases of positive and negative shear should both be considered. As a conservative approach it is assumed that these values apply at all altitudes.
TABLE 1
1 A shear of 0175 sec -1 (104 kt/1000 ft) over a height band of 1000 ft.
2 A shear of 0095 sec -1 (565 kt/1000 ft) over a height band of 3000 ft.
3 A shear of 0065 sec -1 (386 kt/1000 ft) over a height band of 5000 ft.
4 A shear of 005 sec -1 (297 kt/1000 ft) over a height band of 7000 ft.
2.3 The entry of the aeroplane into horizontal and vertical wind shears will be treated as separate cases. Because penetration of these large scale phenomena is fairly slow, recovery action by the pilot is usually possible, and may be beneficial or adverse. In the case of 'manual' flight (i.e. when flight is being controlled by inputs made by the pilot) the aeroplane is assumed to maintain constant attitude until at least 5 seconds after the operation of the excess speed warning device at which time recovery action may be started by use of primary aerodynamic controls alone at a normal acceleration of 15g or the maximum available, whichever is lower. If this recovery action is adverse, it should be assumed that power could be reduced after a further 5 seconds.
The windshear values given in the above table are True Airspeed (TAS). ]
[ ACJ 25.335(d)
Design Speed for Maximum Gust Intensity: (Acceptable Means of Compliance)
See JAR 25.335(d)
In the absence of a more rational analysis the gust load factors may be calculated as follows:

where -
gust alleviation factor;
aeroplane mass ration;
Ude = specified derived gust velocity (fps);
r = density of air (slugs/cu ft);
W/S = wing loading (psf);
= mean geometric chord of the wing (ft);
g = acceleration due to gravity (ft/sec 2)
V = aeroplane's equivalent speed (knots); and
a = slope of the aeroplane normal force coefficient CNA per radian. ]

ACJ 25.337(d)
Limit Manoeuvring Load Factors (Interpretative Material)
See JAR 25.337(d)
Limitations in control movement will not normally be accepted as a sufficient justification for reducing the manoeuvring load factor.

ACJ 25.341(b)
[ Strength and Deformation (Interpretative Material) ]
See JAR 25.341(b)
[ 1 General. When effects of dynamic response to turbulence are assessed by the continuous turbulence method, the following criteria can be used:
2 Continuous Gust Design Criteria.* The gust loads criteria of this paragraph 2 should be applied to mission analysis or design envelope analysis.
2.1 The limit gust loads utilising the continuous turbulence concept should be determined in accordance with the provisions of either paragraph 2.2 or paragraphs 2.3 and 2.4. For components stressed by both vertical and lateral components of turbulence, the resultant combined stress should be considered. The combined stress may be determined on the assumption that vertical and lateral components are uncorrelated.
2.2 Design Envelope Analysis. The limit loads should be determined in accordance with this paragraph 2.2.
2.2.1 All critical altitudes, weights, and weight distributions, as specified in JAR 25.321(b)(1) to (b)(3), and all critical speeds within the ranges indicated in paragraph 2.3, should be considered.]
[ 2.2.2 Values of A (ratio of root-mean-square incremental load to root-mean-square gust velocity) should be determined by dynamic analysis. The power spectral density of the atmospheric turbulence should be as given by the equation-

where-
F = power-spectral density, (ft/s)2 / rad/ft ((m/s)2 / rad/m)
s = true root mean square gust velocity, ft/s (m/s)
W = reduced frequency, rad/ft (rad/m)
L = 2500 ft (762 m)
2.2.3 The limit loads should be obtained by multiplying the values given by the dynamic analysis by the following value of Us (true gust velocity)
a. At speed VC
i. Us = 85 ft/s (2591 m/s) true gust velocity in the interval 0 to 30 000 ft (9144 m) altitude and is linearly decreased to 30 ft/s (914 m/s) true gust velocity at 80 000 ft (24 384 m) altitude.
ii. Us values less than those specified in sub-paragraph a.i. may be used where the applicant can show by rational means that the gust velocity selected is adequate for the aeroplane being considered.
However, the Us values used may not be less than 75 ft/s (2286 m/s), with a linear decrease from that value at 20 000 ft (6096 m) to 30 ft/s (914 m/s) at 80 000 ft (24 384 m).
b. At speed VB. Us is given by 132 times the values obtained under a.
c. At speed VD. Us is given by 05 times the values obtained under a.
d. At speeds between VB and VC, and between VC and VD. Us is given by linear interpolation.
2.2.4 When a stability augmentation system is included in the analysis, the effect of system non-linearities on loads at the limit load level should be realistically or conservatively accounted for.
2.3 Mission analysis. Limit loads should be determined in accordance with this paragraph 2.3.
2.3.1 The expected utilisation of the aeroplane should be represented by one or more flight profiles in which the load distribution and the variation with time of speed, altitude, gross weight, and centre of gravity position are defined. These profiles should be divided into mission segments, or blocks for analysis and average or effective values of the pertinent parameters defined for each segment.
2.3.2 For each of the mission segments defined under paragraph 2.3.1 values of and N0 should be determined by dynamic analysis. is defined as the ratio of root-mean-square incremental load to root-mean-square gust velocity and N0 as the radius of gyration of the load power-spectral density function about zero frequency. The power spectral density of the atmospheric turbulence should be given by the equation in paragraph 2.2.2.
2.3.3 For each of the load and stress quantities selected, the frequency of exceedance should be determined as a function of load level by means of the equation,

where -
y = net values of the load or stress
y one-g = value of the load or stress in one-g level flight
N(y) = average number of exceedances of the indicated values of the load
or stress in unit time
S = symbol denoting summation over all mission segments
t = fraction of total flight time in the given segment
N0, = parameters determined by dynamic analysis as defined in
paragraph 2.3.2
P1 , P2, b1 , b2 = parameters defining the probability distributions of root-mean-square
gust velocity to read from Figures 1 and 2. ]
[ 2.3.4 The limit gust loads should be read from the frequency of exceedance curves at a frequency of exceedance of 2 x 10-5 exceedances per hour. Both positive and negative load directions should be considered in determination of the limit loads.
2.3.5 If a stability augmentation system is utilised to reduce the gust loads, consideration should be given to the fraction of flight time that the system may be inoperative. The flight profiles of paragraph 2.3.1 should include flight with the system inoperative for this fraction of the flight time. When a stability augmentation system is included in the analysis, the effect of system non-linearities on loads at the limit load level should be realistically or conservatively accounted for.
2.4 Supplementary Design Envelope Analysis. In addition to the limit loads defined by paragraph 2.3, limit loads should also be determined in accordance with paragraph 2.2, modified as follows:-
2.4.1 In paragraph 2.2.3 a. the values of Us = 85 ft/s (2591 m/s) true is replaced by Us = 60 ft/s (1829 m/s) true in the interval 0 to 30 000 ft (9144 m) altitude and is linearly decreased to 25 ft/s (762 m/s) true at 80 000 ft (24 384 m) altitude.
2.4.2 In paragraph 2.2, the reference to paragraphs 2.2.3a. to 2.2.3c. should be understood as referring to the paragraph as modified by paragraph 2.4.

FIGURE 1 - P1 AND P2 VALUES ]



ACJ.25.345(a)
High Lift Devices (Gust Conditions) (Acceptable Means of Compliance)
See JAR 25.345(a)
Compliance with JAR 25.345(a) may be demonstrated by an analysis in which the solution of the vertical response equations is made by assuming the aircraft to be rigid. If desired, the analysis may take account of the effects of structural flexibility on a quasi-flexible basis (i.e. using aerodynamic derivatives and load distributions corresponding to the distorted structure under maximum gust load). ]

ACJ 25.345(c)
High Lift Devices (Procedure Flight Condition) (Interpretative Material)
See JAR 25.345(c)
1 En-route conditions are flight segments other than take-off, approach and landing. As applied to the use of high lift devices the following flight phases are to be included in en-route conditions:
- holding in designated areas outside the terminal area of the airport, and
- flight with flaps extended from top of descent.
The following flight phases are not to be included in en-route conditions:
- portion of the flight corresponding to standard arrival routes preceding the interception of the final approach path, and
- holding at relatively low altitude close to the airport.
[ 2 To apply JAR 25.341 (a) gust conditions to JAR 25.345(c), the speeds VFC and VFD should be determined for the flap positions selected in en-route conditions. ]
These procedures should ensure proper speed margins for flap retraction in the case of severe turbulence when the aeroplane is in a low speed en-route holding configuration.
3 The manoeuvre of JAR 25.345(c)(1) is to be considered as a balanced condition. (See JAR 25.331(b) for definition.)
[ ACJ 25.349(b)
Unsymmetrical Gusts (Acceptable Means of Compliance)
See JAR 25.349(b)
In the absence of a more rational analysis, the maximum air load due to gust (Lg), applied to the aircraft may be computed as follows:
Lg =
where-
Kg = = gust alleviation factor;
mg = = aeroplane mass ratio;
Ude = specified derived gust velocity (fps);
r = density of air (slugs/cu ft);
W/S = wing loading (psf);
= mean geometric chord of the wing (ft);
g = acceleration due to gravity (ft/sec2);
V = aeroplane's equivalent speed (knots);
a = slope of the aeroplane normal force coefficient CNA per radian; and
S = wing area (ft2).
For the purposes of this requirement the wing air load should be calculated from the maximum air load due to gust by using a representative distribution of load amongst the major lifting surfaces. ]
[ ACJ 25.365(e)
Pressurised Compartment Loads (Interpretative Material)
See JAR 25.365(e)
The computed opening size from 25.365(e)(2) should be considered only as a mathematical means of developing ultimate pressure design loads to prevent secondary structural failures. No consideration need be given to the actual shape of the opening, nor to its exact location on the pressure barrier in the compartment. The damage and loss of strength at the opening location should not be considered.
A hazard assessment should determine which structures should be required to withstand the resulting differential pressure loads. The assessment of the secondary consequences of failures of these structures should address those events that have a reasonable probability of interfering with safe flight and landing, for example failures of structures supporting critical systems. For this assessment the risk of impact on the main structure from non critical structures, such as fairings, detached from the aircraft due to decompression need not be considered. ]

ACJ 25.393(a)
Loads Parallel to Hinge Line (Interpretative Material)
See JAR 25.393(a)
The loads parallel to the hinge line on primary control surfaces and other movable surfaces, such as tabs, spoilers, speedbrakes, flaps, slats and all-moving tailplanes, should take account of axial play between the surface and its supporting structure in complying with JAR 25.393(a). For the rational analysis, the critical airframe acceleration time history in the direction of the hinge line from all flight and ground design conditions (except the emergency landing conditions of JAR 25.561) should be considered. The play assumed in the control surface supporting structure, should include the maximum tolerable nominal play and the effects of wear. In showing compliance with this paragraph it would be acceptable to assume a linear mass-spring system with no damping.
[ ACJ 25.427(b)(1)
Unsymmetrical Loads (Acceptable Means of Compliance)
See JAR 25.427(b)(1)
For the purpose of calculating unsymmetric loads from symmetric vertical gust conditions and in the absence of a more rational analysis using loads calculated in accordance with JAR 25.341(a), the limit symmetric vertical gust loads on the horizontal tail should be determined in accordance with the provisions of JAR 25.349(b)(1) and (b)(2). ]

ACJ 25.427(b)(2)
Unsymmetrical Loads (Interpretative Material)
See JAR 25.427(b)(2)
The following wording expresses more clearly the intent of JAR 25.427(b)(2):
The empennage arrangements where the horizontal tail surfaces have appreciable dihedral or are supported by the vertical tail surfaces, the surfaces and supporting structure must be designed for the prescribed flight load conditions. For each condition, considered separately, the resultant loads on the vertical and horizontal surfaces must be combined for stressing purposes.

ACJ 25.491
Take-off Run (Acceptable Means of Compliance)
See JAR 25.491
In the absence of a more rational analysis the following can be considered as an Acceptable Means of Compliance:
a. At attitudes ranging from JAR 25.479(e)(1) to JAR 25.481(c) for the condition of combined vertical, side and drag loads, a drag and side load of 20% respectively, of the ground reaction should be combined with this ground reaction where the latter is defined as 150% of WT/2 on each main gear, where WT is the design take-off weight of the aeroplane.
b. With the nose and main gear in contact with the ground the condition of a vertical load equal to 17 times the static ground reaction should be investigated under the most adverse aeroplane loading distribution at design take-off weight, taking into account thrust from the engines.

ACJ 25.493(c)
Braked Roll Conditions (Interpretative Material)
See JAR 25.493(c)
1 In seeking a reduced drag reaction for the total aeroplane, the most likely consideration would be limitations in brake energy absorption capability. The drag due to this would be derived from the maximum value of the summation of per wheel fitted with brakes, where -
T = the time dependent brake torque the maximum value of which corresponds
to the Certified Maximum Braking Torque T1, and
r = the rolling radius under normal tyre pressure and appropriate vertical reaction.
In the absence of a more rational determination, T may be made equal to T1.
2 The Certified Maximum Braking Torque for each wheel fitted with brakes will be determined as the value never likely to be exceeded during the operation of the aeroplane. This will be equal to the product of the maximum recorded brake torque and the production variability factor. The maximum recorded brake torque will be established from tests covering all practical ranges of brake operating conditions likely to be encountered. In particular, the ranges of speed, temperature and operating pressure of the brake should be considered. In the absence of better evidence the production variability factor, used in the determination of T and T1, may be taken as 133.

ACJ 25.493(d)
Braked Roll Conditions (Acceptable Means of Compliance)
See JAR 25.493(d)
1 For calculation of a coefficient of friction corresponding to the maximum main gear braking effort it is permissible to take into account the effects of any anti-skid system and a minimum tyre slip ratio appropriate to the most critical taxi speed. Unless a lower value can be substantiated the drag reaction at each main wheel fitted with brakes should be taken as equal to the vertical reaction multiplied by a coefficient of friction of 08.
[2 In the absence of a more rational analysis, the nose gear vertical reaction may be determined as follows:

Where: VN = nose gear vertical reaction.
W = maximum take-off weight.
A = distance between the c.g. of the aeroplane and the
nose wheel, measured horizontally.
B = distance between the c.g. of the aeroplane and the line
joining the centres of the main wheels horizontally.
E = vertical height of the c.g. of the aeroplane above
ground in the 1g condition.*
= coefficient of friction.
f = dynamic response factor; 20 is to be used unless
a lower factor is substantiated (See (3) below).
*See also JAR-25 Section 1, Appendix A, Figure 1
3 In the absence of other information, the dynamic response factor f may be defined by the equation:
f = 1 + exp
where x is the effective critical damping ratio of the rigid body pitching mode. ]
[ ACJ 25.561
General (Acceptable Means of Compliance)
See JAR 25.561
In complying with the provisions of JAR 25.561(b) & (c), the loads arising from the restraint of seats and items of equipment etc. should be taken into the structure to a point where the stresses can be dissipated (e.g. for items attached to the fuselage floor, the load paths from the attachments through to the fuselage primary structure should be taken into account). ]

ACJ 25.561(c)
General (Interpretative Material)
See JAR 25.561(c)
[ The local attachments subject to severe wear and tear refer to the attachments of these items which might be frequently removed and installed (such as seats, attachment, belt and harness attachments, trolley, carts retainers etc.). ]
[ ACJ 25.561(d)
General (Acceptable Means of Compliance)
See JAR 25.561(d)
For the local attachments of seats and items of mass it should be shown by analysis and/or tests that under the specified load conditions, the intended retaining function in each direction is still available. ]

ACJ 25.571(a)
Damage Tolerance and Fatigue Evaluation of Structure (Acceptable Means of Compliance)
See JAR 25.571(a) and JAR 25.571(e)
1 Introduction
1.1 The contents of this ACJ are considered by the Authority in determining compliance with the damage-tolerance and fatigue requirements of JAR 25.571.
1.1.1 Although a uniform approach to the evaluation required by JAR 25.571 is desirable, it is recognised that in such a complex field new design features and methods of fabrication, new approaches to the evaluation, and new configurations could necessitate variations and deviations from the procedures described in this ACJ.
1.1.2 Damage-tolerance design is required, unless it entails such complications that an effective damage-tolerant structure cannot be achieved within the limitations of geometry, inspectability, or good design practice. Under these circumstances, a design that complies with the fatigue evaluation (safe-life) requirements is used. Typical examples of structure that might not be conducive to damage-tolerance design are landing gear, engine mounts, and their attachments.
1.1.3 Experience with the application of methods of fatigue evaluation indicate that a test background should exist in order to achieve the design objective. Even under the damage-tolerance method discussed in paragraph 2, 'Damage-tolerance (fail-safe) evaluation', it is the general practice within industry to conduct damage-tolerance tests for design information and guidance purposes. Damage location and growth data should also be considered in establishing a recommended inspection programme.
1.1.4 Assessing the fatigue characteristics of certain structural elements, such as major fittings, joints, typical skin units, and splices, to ensure that the anticipated service life can reasonably be attained, is needed for structure to be evaluated under JAR 25.571(c).
1.2 Typical Loading Spectra Expected in Service. The loading spectrum should be based on measured statistical data of the type derived from government and industry load history studies and, where insufficient data are available, on a conservative estimate of the anticipated use of the aeroplane. The principal loads that should be considered in establishing a loading spectrum are flight loads (gust and manoeuvre), ground loads (taxiing, landing impact, turning, engine runup, braking, and towing) and pressurisation loads. The development of the loading spectrum includes the definition of the expected flight plan which involves climb, cruise, descent, flight times, operational speeds and altitudes, and the approximate time to be spent in each of the operating regimes. Operations for crew training, and other pertinent factors, such as the dynamic stress characteristics of any flexible structure excited by turbulence, should also be considered. For pressurised cabins, the loading spectrum should include the repeated application of the normal operating differential pressure, and the super-imposed effects of flight loads and external aerodynamic pressures.
1.3 Components to be Evaluated. In assessing the possibility of serious fatigue failures, the design should be examined to determine probable points of failure in service. In this examination, consideration should be given, as necessary, to the results of stress analyses, static tests, fatigue tests, strain gauge surveys, tests of similar structural configurations, and service experience. Service experience has shown that special attention should be focused on the design details of important discontinuities, main attachment fittings, tension joints, splices, and cutouts such as windows, doors and other openings. Locations prone to accidental damage (such as that due to impact with ground servicing equipment near aeroplane doors) or to corrosion should also be considered.
1.4 Analyses and Tests. Unless it is determined from the foregoing examination that the normal operating stresses in specific regions of the structure are of such a low order that serious damage growth is extremely improbable, repeated load analyses or tests should be conducted on structure representative of components or sub-components of the wing, control surfaces, empennage, fuselage, landing gear, and their related primary attachments. Test specimens should include structure representative of attachment fittings, major joints, changes in section, cutouts, and discontinuities. Any method used in the analyses should be supported, as necessary, by test or service experience. Generally it will be required to substantiate the primary structure against the provisions of JAR 25.571(b) and (c) by representative testing. The nature and extent of tests on complete structures or on portions of the primary structure will depend upon applicable previous design and structural tests, and service experience with similar structures. The scope of the analyses and supporting test programmes should be agreed with the Authority.
1.5 Repeated Load Testing. In the event of any repeated load testing necessary to support the damage tolerance or safe-life objectives of JAR 25.571(b) and (c) respectively not being concluded at the issuance of type certificate, at least one year of safe operation should be substantiated at the time of certification. In order not to invalidate the certificate of airworthiness the fatigue substantiation should stay sufficiently ahead of the service exposure of the lead aeroplane.
2 Damage-tolerance (Fail-safe) Evaluation
2.1 General. The damage-tolerance evaluation of structure is intended to ensure that should serious fatigue, corrosion, or accidental damage occur within the operational life of the aeroplane, the remaining structure can withstand reasonable loads without failure or excessive structural deformation until the damage is detected. Included are the considerations historically associated with fail-safe design. The evaluation should encompass establishing the components which are to be designed as damage-tolerant, defining the loading conditions and extent of damage, conducting sufficient representative tests and/or analyses to substantiate the design objectives (such as life to crack-initiation, crack propagation rate and residual strength) have been achieved and establishing data for inspection programmes to ensure detection of damage. Interpretation of the test results should take into account the scatter in crack propagation rates as well as in lives to crack-initiation. Test results should be corrected to allow for variations between the specimen and the aeroplane component thickness and sizes. This evaluation applies to either single or multiple load path structure.
2.1.1 Design features which should be considered in attaining a damage-tolerant structure include the following:
a. Multiple load path construction and the use of crack stoppers to control the rate of crack growth, and to provide adequate residual static strength;
b. Materials and stress levels that, after initiation of cracks, provide a controlled slow rate of crack propagation combined with high residual strength. For single load path discrete items, such as control surface hinges, wing spar joints or stabiliser pivot fittings the failure of which could be catastrophic, it should be clearly demonstrated that cracks starting from material flaws, manufacturing errors or accidental damage (including corrosion) have been properly accounted for in the crack propagation estimate and inspection method;
c. Arrangement of design details to ensure a sufficiently high probability that a failure in any critical structural element will be detected before the strength has been reduced below the level necessary to withstand the loading conditions specified in JAR 25.571(b) so as to allow replacement or repair of the failed elements; and
d. Provisions to limit the probability of concurrent multiple damage, particularly after long service, which could conceivably contribute to a common fracture path. The achievement of this would be facilitated by ensuring sufficient life to crack-initiation. Examples of such multiple damage are -
i. A number of small cracks which might coalesce to form a single long crack;
ii. Failures, or partial failures, in adjacent areas, due to the redistribution of loading following a failure of a single element; and
iii. Simultaneous failure, or partial failure, of multiple load path discrete elements, working at similar stress levels.
In practice it may not be possible to guard against the effects of multiple damage and fail-safe substantiation may be valid only up to a particular life which would preclude multiple damage.
e. The aeroplane may function safely with an element missing. This feature would be admitted only, provided its separation will not prevent continued safe flight and landing and the probability of occurrence is acceptably low.
2.1.2 In the case of damage which is readily detectable within a short period (50 flights, say) for which JAR 25.571(b) allows smaller loads to be used, this relates to damage which is large enough to be detected by obvious visual indications during walk around, or by indirect means such as cabin pressure loss, cabin noise, or fuel leakage. In such instances, and in the absence of a probability approach the residual load levels except for the trailing edge flaps may be reduced to not less than the following:
a. The maximum normal operating differential pressure (including the expected external aerodynamic pressures under 1g level flight) multiplied by a factor of 110 omitting other loads.
b. 85% of the limit flight manoeuvre and ground conditions of JAR 25.571(b)(1) to (6) inclusive, excluding (5)(ii) and separately 75% of the limit gust velocities (vertical or lateral) as specified at speeds up to VC in JAR 25.571(b)(2) and (b)(5)(i). On the other hand if the probability approach is used the residual load levels may not in any case be lower than the values given in paragraph 2.7.2 of this ACJ for one flight exposure. In the case where fatigue damage is arrested at a readily detectable size following rapid crack growth or a sudden load path failure under the application of high loads, the structure must be able to withstand the loads defined in JAR 25.571(b)(1) to (6) inclusive up to that size of damage. For the subsequent growth of that damage, lower loads as stated above may be used.
2.2 Identification of Principal Structural Elements. Principal structural elements are those which contribute significantly to carrying flight, ground, and pressurisation loads, and whose failure could result in catastrophic failure of the aeroplane. Typical examples of such elements are as follows:
2.2.1 Wing and empennage
a. Control surfaces, slats, flaps and their attachment hinges and fittings;
b. Integrally stiffened plates;
c. Primary fittings;
d. Principal splices;
e. Skin or reinforcement around cutouts or discontinuities;
f. Skin-stringer combinations;
g. Spar caps; and
h. Spar webs.
2.2.2 Fuselage
a. Circumferential frames and adjacent skin;
b. Door frames;
c. Pilot window posts;
d. Pressure bulkheads;
e. Skin and any single frame or stiffener element around a cutout;
f. Skin or skin splices, or both, under circumferential loads;
g. Skin or skin splices, or both, under fore-and-aft loads;
h. Skin around a cutout;
i. Skin and stiffener combinations under fore-and-aft loads; and
j. Window frames.
2.3 Extent of Damage. Each particular design should be assessed to establish appropriate damage criteria in relation to inspectability and damage-extension characteristics. In any damage determination, including those involving multiple cracks, it is possible to establish the extent of damage in terms of detectability with the inspection techniques to be used, the associated initially detectable crack size, the residual strength capabilities of the structure, and the likely damage-extension rate considering the expected stress redistribution under the repeated loads expected in service and with the expected inspection frequency. Thus, an obvious partial failure could be considered to be the extent of the damage or residual strength assessment, provided a positive determination is made that the fatigue cracks will be detectable by the available inspection techniques at a sufficiently early stage of the crack development. In a pressurised fuselage, an obvious partial failure might be detectable through the inability of the cabin to maintain operating pressure or controlled decompression after occurrence of the damage. The following are typical examples of partial failures which should be considered in the evaluation:
2.3.1 Detectable skin cracks emanating from the edge of structural openings or cutouts;
2.3.2 A detectable circumferential or longitudinal skin crack in the basic fuselage structure;
2.3.3 Complete severence of interior frame elements or stiffeners in addition to a detectable crack in the adjacent skin;
2.3.4 A detectable failure of one element where dual construction is utilised in components such as spar caps, window posts, window or door frames, and skin structure;
2.3.5 The presence of a detectable fatigue failure in at least the tension portion of the spar web or similar element; and
2.3.6 The detectable failure of a primary attachment, including a control surface hinge and fitting.
2.4 Inaccessible Areas. Every reasonable effort should be made to ensure inspectability of all structural parts, and to qualify them under the damage-tolerance provisions. In those cases where inaccessible and uninspectable blind areas exist, and suitable damage tolerance cannot practically be provided to allow for extension of damage into detectable areas, the structure should be shown to comply with the fatigue (safe-life) requirements in order to ensure its continued airworthiness. In this respect particular attention should be given to the effects of corrosion.
2.5 Testing of Principal Structural Elements. The nature and extent of tests on complete structures or on portions of the primary structure will depend upon applicable previous design, construction, tests, and service experience, in connection with similar structures. Simulated cracks should be as representative as possible of actual fatigue damage. Where it is not practical to produce actual fatigue cracks, damage can be simulated by cuts made with a fine saw, sharp blade, guillotine, or other suitable means. In those cases where bolt failure, or its equivalent, is to be simulated as part of a possible damage configuration in joints or fittings, bolts can be removed to provide that part of the simulation, if this condition would be representative of an actual failure under typical load. Where accelerated crack propagation tests are made, the possibility of creep cracking under real time pressure conditions should be recognised especially as the crack approaches its critical length.
2.6 Identification of Locations to be Evaluated. The locations of damage to structure for damage-tolerances evaluation should be identified as follows:
2.6.1 Determination of General Damage Locations. The location and modes of damage can be determined by analysis or by fatigue tests on complete structures or subcomponents. However, tests might be necessary when the basis for analytical prediction is not reliable, such as for complex components. If less than the complete structure is tested, care should be taken to ensure that the internal loads and boundary conditions are valid. Any tests should be continued sufficiently beyond the expected service life to ensure that, as far as practicable, the likely locations and extent of crack initiation are discovered.
a. If a determination is made by analysis, factors such as the following should be taken into account:
i. Strain data on undamaged structure to establish points of high stress concentration as well as the magnitude of the concentration;
ii. Locations where permanent deformation occurred in static tests;
iii. Locations of potential fatigue damage identified by fatigue analysis; and
iv. Design details which service experience of similarly designed components indicate are prone to fatigue or other damage.
b. In addition, the areas of probable damage from sources such as corrosion, disbonding, accidental damage or manufacturing defects should be determined from a review of the design and past service experience.
2.6.2 Selection of Critical Damage Areas. The process of actually locating where damage should be simulated in principal structural elements identified in paragraph 2.2 of this ACJ should take into account factors such as the following:
a. Review analysis to locate areas of maximum stress and low margin of safety;
b. Selecting locations in an element where the stresses in adjacent elements would be the maximum with the damage present;
c. Selecting partial fracture locations in an element where high stress concentrations are present in the residual structure; and
d. Selecting locations where detection would be difficult.
2.7 Damage-tolerance Analysis and Tests. It should be determined by analysis, supported by test evidence, that the structure with the extent of damage established for residual strength evaluation can withstand the specified design limit loads (considered as ultimate loads), and that the damage growth rate under the repeated loads expected in service (between the time at which the damage becomes initially detectable and the time at which the extent of damage reaches the value for residual strength evaluation) provides a practical basis for development of the inspection programme and procedures described in paragraph 2.8 of this ACJ. The repeated loads should be as defined in the loading, temperature, and humidity spectra. The loading conditions should take into account the effects of structural flexibility and rate of loading where they are significant.
2.7.1 The damage-tolerance characteristics can be shown analytically by reliable or conservative methods such as the following:
a. By demonstrating quantitative relationships with structure already verified as damage tolerant;
b. By demonstrating that the damage would be detected before it reaches the value for residual strength evaluation; or
c. By demonstrating that the repeated loads and limit load stresses do not exceed those of previously verified designs of similar configuration, materials and inspectability.
2.7.2 The maximum extent of immediately obvious damage from discrete sources should be determined and the remaining structure shown to have static strength for the maximum load (considered as ultimate load) expected during the completion of the flight. In the absence of a rational analysis the following ultimate loading conditions should be covered:
a. At the time of the incident:
i. The maximum normal operating differential pressure (including the expected external aerodynamic pressures during 1g level flight) multiplied by a factor 11 combined with 1g flight loads.
ii. The aeroplane, assumed to be in 1g level flight should be shown to be able to survive the overswing condition due to engine thrust asymmetry and pilot corrective action taking into account any damage to the flight controls which it is presumed the aeroplane has survived.
b. Following the incident: 70% limit flight manoeuvre loads and, separately, 40% of the limit gust velocity (vertical or lateral) as specified at VC up to the maximum likely operational speed following failure, each combined with the maximum appropriate cabin differential pressure (including the expected external aerodynamic pressures). Further, any loss in structural stiffness which might arise shall be shown to result in no dangerous reduction in freedom from flutter up to speed VC/MC.
2.8 Inspection. Detection of damage before it becomes dangerous is the ultimate control in ensuring the damage-tolerance characteristics of the structure. Therefore, the applicant should provide sufficient guidance information to assist operators in establishing the frequency, extent, and methods of inspection of the critical structure, and this kind of information must, under JAR 25.571(a)(3), be included in the maintenance manual required by JAR 25.1529. Due to the inherent complex interactions of the many parameters affecting damage tolerance, such as operating practices, environmental effects, load sequence on crack growth, and variations in inspection methods, related operational experience should be taken into account in establishing inspection procedures. It is extremely important to ensure by regular inspection the detection of damage in areas vulnerable to corrosion or accidental damage. However for crack initiation arising from fatigue alone, the frequency and extent of the inspections may be reduced during the period up to the demonstrated crack-free life of the part of the structure, including appropriate scatter factors (see paragraph 3.2). Comparative analysis can be used to guide the changes from successful past practice when necessary. Therefore, maintenance and inspection requirements should recognise the dependence on experience and should be specified in a document that provides for revision as a result of operational experience, such as the one containing the Manufacturers Recommended Structural Inspection Programme.
3 Fatigue (Safe-Life) Evaluation
3.1 General. The evaluation of structure under the following fatigue (safe-life) strength evaluation methods is intended to ensure that catastrophic fatigue failure, as a result of the repeated loads of variable magnitude expected in service, is extremely improbable throughout the structure's operational life. Under these methods, loading spectra should be established, the fatigue life of the structure for the spectra should be determined, and a scatter factor should be applied to the fatigue life to establish the safe-life for the structure. The evaluation should include the following; however, in some instances it might be necessary to correlate the loadings used in the analysis with flight load and strain surveys:
3.1.1 Estimating or measuring the expected loading spectra for the structure;
3.1.2 Conducting a structural analysis including consideration of the stress concentration effects;
3.1.3 Fatigue testing of structure which cannot be related to a test background to establish response to the typical loading spectrum expected in service;
3.1.4 Determining reliable replacement times by interpreting the loading history, variable load analyses, fatigue test data, service experience, and fatigue analyses; and
3.1.5 Providing data for inspection and maintenance instructions and guidance information to the operators.
3.1.6 In addition fatigue initiation from sources such as corrosion, stress corrosion, disbonding, accidental damage and manufacturing defects should be covered based on a review of the design and past service experience.
3.2 Safe-Life Determinations: Scatter Factor. In the interpretation of fatigue analyses and test data, the effect of variability should, under JAR 25.571(c), be accounted for by an appropriate scatter factor. There are a number of considerations peculiar to each design and test that necessitate evaluation by the applicant. These considerations will depend on the scope of the analyses and supporting test evidence and for example the number and representativeness of test specimens, the material, the type of repeated load test, the extent of information on the expected loading spectra, consequence of failure and environmental conditions.
3.3 Replacement Times. Replacement times should be established for parts with established safe-lives and should, under JAR 25.571(a)(3), be included in the information prepared under JAR 25.1529. These replacement times can be extended if additional data indicates an extension is warranted. Important factors which should be considered for such extensions include, but are not limited to, the following:
3.3.1 Comparison of original evaluation with service experience;
3.3.2 Recorded Load and Stress Data. Recorded load and stress data entails instrumenting aeroplanes in service to obtain a representative sampling of actual loads and stresses experienced. The data to be measured includes airspeed, altitude, and load factor versus time data; or airspeed, altitude and strain ranges versus time data; or similar data. This data, obtained by instrumenting aeroplanes in service, provides a basis for correlating the estimated loading spectrum with the actual service experience;
3.3.3 Additional Analyses and Tests. If test data and analyses based on repeated load tests of additional specimens are obtained, a re-evaluation of the established safe-life can be made;
3.3.4 Tests of Parts Removed from Service. Repeated load tests of replaced parts can be utilised to re-evaluate the established safe-life. The tests should closely simulate service loading conditions. Repeated load testing of parts removed from service is especially useful where recorded load data obtained in service are available since the actual loading experienced by the part prior to replacement is known; and
3.3.5 Repair or Rework of the Structure. In some cases, repair or rework of the structure can gain further life.
3.4 Type Design Developments and Changes. For design developments, or design changes, involving structural configurations similar to those of a design already shown to comply with the applicable provisions of JAR 25.571(c), it might be possible to evaluate the variations in critical portions of the structure on a comparative basis. Typical examples would be redesign of the wing structure for increased loads, and the introduction in pressurised cabins of cutouts having different locations or different shapes, or both. This evaluation should involve analysis of the predicted stresses of the redesigned primary structure and correlation of the analysis with the analytical and test results used in showing compliance of the original design with JAR 25.571(c).

ACJ 25.571(b)
Damage-tolerance (fail-safe) Evaluation (Interpretative Material)
See JAR 25.571(b) and JAR 25.571(e)
In the above mentioned conditions the dynamic effects are included except that if significant changes in stiffness and/or geometry follow from the failure or partial failure the response should be further investigated.

ACJ 25.581
Lightning Protection (Acceptable Means of Compliance and Interpretative Material)
See JAR 25.581
1 External Metal Parts
1.1 External metal parts should either be -
a. Electrically bonded to the main earth system by primary bonding paths, or
b. So designed and/or protected that a lightning discharge to the part (e.g. a radio aerial) will cause only local damage which will not endanger the aeroplane or its occupants.
1.2 In addition, where internal linkages are connected to external parts (e.g. control surfaces), the linkages should be bonded to main earth or airframe by primary bonding paths as close to the external part as possible.
1.3 Where a primary conductor provides or supplements the primary bonding path across an operating jack (e.g. on control surfaces or nose droop) it should be of such an impedance and so designed as to limit to a safe value the passage of current through the jack.
1.4 In considering external metal parts, consideration should be given to all flight configurations (e.g. lowering of landing gear and wing-flaps) and also the possibility of damage to the aeroplane electrical system due to surges caused by strikes to protuberances (such as pitot heads) which have connections into the electrical system.
2 External Non-metallic Parts
2.1 External non-metallic parts should be so designed and installed that -
a. They are provided with effective lightning diverters which will safely carry the lightning discharges described in Table 1 of ACJ 25X899,
b. Damage to them by lightning discharges will not endanger the aeroplane or its occupants, or
c. A lightning strike on the insulated portion is improbable because of the shielding afforded by other portions of the aeroplane.
Where lightning diverters are used the surge carrying capacity and mechanical robustness of associated conductors should be at least equal to that required for primary conductors.
2.2 Where unprotected non-metallic parts are fitted externally to the aeroplane in situations where they may be exposed to lightning discharges (e.g. radomes) the risks include the following:
a. The disruption of the materials because of rapid expansion of gases within them (e.g. water vapour),
b. The rapid build up of pressure in the enclosures provided by the parts, resulting in mechanical disruption of the parts themselves or of the structure enclosed by them,
c. Fire caused by the ignition of the materials themselves or of the materials contained within the enclosures, and
d. Holes in the non-metallic part which may present a hazard at high speeds.
2.3 The materials used should not absorb water and should be of high dielectric strength in order to encourage surface flash-over rather than puncture. Laminates made entirely from solid material are preferable to those incorporating laminations of cellular material.
2.4 Those external non-metallic part which is not classified as primary structure should be protected by primary conductors.
2.5 Where damage to an external non-metallic part which is not classified as primary structure may endanger the aeroplane, the part should be protected by adequate lightning diverters.
2.6 Confirmatory tests may be required to check the adequacy of the lightning protection provided (e.g. to confirm the adequacy of the location and size of bonding strips on a large radome.)
ACJ - Subpart D

ACJ 25.603
Composite Aircraft Structure (Acceptable Means of Compliance)
See JAR 25.603
1 Purpose. This ACJ sets forth an acceptable means, but not the only means, of showing compliance with the provisions of JAR-25 regarding airworthiness type certification requirements for composite aircraft structures, involving fibre-reinforced materials, e.g. carbon (graphite), boron, aramid (Kevlar), and glass-reinforced plastics. Guidance information is also presented on associated quality control and repair aspects.
This ACJ material is identical, apart from minor editing, to the structural content of FAA Advisory Circular AC 20.107A, dated 25 April 1984.
The individual JAR paragraphs applicable to each ACJ paragraph are listed in Table 1 of this ACJ.
2 Definitions
2.1 Design values. Material, structural element, and structural detail properties that have been determined from test data and chosen to assure a high degree of confidence in the integrity of the completed structure (see JAR 25.613(b)).
2.2 Allowables. Material values that are determined from test data at the laminate or lamina level on a [ probability basis (e.g. A or B base values). ]
2.3 Laminate level design values or allowables. Established from multi-ply laminate test data and/or from test data at the lamina level and then established at the laminate level by test validated analytical methods.
2.4 Lamina level material properties. Established from test data for a single-ply or multi-ply single-direction oriented lamina lay-up.
2.5 Point design. An element or detail of a specific design which is not considered generically applicable to other structure for the purpose of substantiation (e.g. lugs and major joints). Such a design element or detail can be qualified by test or by a combination of test and analysis.
2.6 Environment. External, non-accidental conditions (excluding mechanical loading), separately or in combination, that can be expected in service and which may affect the structure (e.g. temperature, moisture, UV radiation, and fuel).
2.7 Degradation. The alteration of material properties (e.g. strength, modulus, coefficient of expansion) which may result from deviations in manufacturing or from repeated loading and/or environmental exposure.
2.8 Discrepancy. A manufacturing anomaly allowed and detected by the planned inspection procedure. They can be created by processing, fabrication or assembly procedures.
2.9 Flaw. A manufacturing anomaly created by processing, fabrication or assembly procedures.
2.10 Damage. A structural anomaly caused by manufacturing (processing, fabrication, assembly or handling) or service usage. Usually caused by trimming, fastener installation or foreign object contact.
2.11 Impact damage. A structural anomaly created by foreign object impact.
2.12 Coupon. A small test specimen (e.g. usually a flat laminate) for evaluation of basic lamina or laminate properties or properties of generic structural features (e.g. bonded or mechanically fastened joints).
2.13 Element. A generic element of a more complex structural member (e.g. skin, stringers, shear panels, sandwich panels, joints, or splices).
[ 2.14 Detail. A non-generic structural element of a more complex member (e.g. specific design configured joints, splices, stringers, stringer runouts, or major access holes).
2.15 Subcomponent. A major three-dimensional structure which can provide complete structural representation of a section of the full structure (e.g. stub-box, section of a spar, wing panel, wing rib, body panel, or frames).
2.16 Component. A major section of the airframe structure (e.g. wing, body, fin, horizontal stabiliser) which can be tested as a complete unit to qualify the structure.
3 General
3.1 This ACJ is published to aid the evaluation of certification programmes for composite applications and reflects the current status of composite technology. It is expected that this ACJ will be modified periodically to reflect technology advances.
3.2 The extent of testing and /or analysis and the degree of environmental accountability required will differ for each structure depending upon the expected service usage, the material selected, the design margins, the failure criteria, the data base and experience with similar structures, and on other factors affecting a particular structure. It is expected that these factors will be considered when interpreting this ACJ for use on a specific application.
4 Material and Fabrication Development
4.1 To provide an adequate design data base, environmental effects on the design properties of the material system should be established.
4.2 Environmental design criteria should be developed that identify the most critical environmental exposures, including humidity and temperature, to which the material in the application under evaluation may be exposed. This is not required where existing data demonstrate that no significant environmental effects, including the effects of temperature and moisture, exist for material systems and construction details, within the bounds of environmental exposure being considered. Experimental evidence should be provided to demonstrate that the material design values or allowables are attained with a high degree of confidence in the appropriate critical environmental exposures to be expected in service. The effect of the service environment on static strength, fatigue and stiffness properties should be determined for the material system through tests (e.g. accelerated environmental tests, or from applicable service data). The effects of environmental cycling (i.e. moisture and temperature) should be evaluated. Existing test data may be used where it can be shown directly applicable to the material system.
4.3 The material system design values or allowables should be established on the laminate level by either test of the laminate or by test of the lamina in conjunction with a test-validated analytical method.
4.4 For a specific structural configuration of an individual component (point design), design values may be established which include the effects of appropriate design features (holes, joints, etc.).
4.5 Impact damage is generally accommodated by limiting the design strain level.
5 Proof of Structure - Static
5.1 The static strength of the composite design should be demonstrated through a programme of component ultimate load tests in the appropriate environment, unless experience with similar designs, material systems and loadings is available to demonstrate the adequacy of the analysis supported by subcomponent tests, or component tests to agreed lower levels. ]
[ 5.2 The effects of repeated loading and environmental exposure which may result in material property degradation should be addressed in the static strength evaluation. This can be shown by analysis supported by test evidence, by tests at the coupon, element or subcomponent level, or alternatively by relevant existing data.
5.3 Static strength structural substantiation tests should be conducted on new structure unless the critical load conditions are associated with structure that has been subjected to repeated loading and environmental exposure. In this case either -
a. The static test should be conducted on structure with prior repeated loading and environmental exposure, or
b. Coupon/Element/Subcomponent test data should be provided to assess the possible degradation of static strength after application of repeated loading and environmental exposure, and this degradation accounted for in the static test or in the analysis of the results of the static test of the new structure.
5.4 The component static test may be performed in an ambient atmosphere if the effects of the environment are reliably predicted by subcomponent and/or coupon tests and are accounted for in the static test or in the analysis of the results of the static test.
5.5 The static test articles should be fabricated and assembled in accordance with production specifications and processes so that the test articles are representative of production structure.
5.6 When the material and processing variability of the composite structure is greater than the variability of current metallic structures, the difference should be considered in the static strength substantiation by -
a. Deriving proper allowables or design values for use in the analysis, and the analysis of the results of supporting tests, or
b. Accounting for it in the static test when static proof of structure is accomplished by component test.
5.7 Composite structures that have high static margins of safety may be substantiated by analysis supported by subcomponent, element and/or coupon testing.
5.8 It should be shown that impact damage that can be realistically expected from manufacturing and service, but not more than the established threshold of detectability for the selected inspection procedure, will not reduce the structural strength below ultimate load capability. This can be shown by analysis supported by test evidence, or by tests at the coupon, element or subcomponent level.
6 Proof of Structure - Fatigue/Damage Tolerance
6.1 The evaluation of composite structure should be based on the applicable requirements of JAR 25.571. The nature and extent of analysis or tests on complete structures and/or portions of the primary structure will depend upon applicable previous fatigue/damage tolerant designs, construction, tests, and service experience on similar structures. In the absence of experience with similar designs, approved structural development tests of components, subcomponents, and elements should be performed. The following considerations are unique to the use of composite material systems and should be observed for the method of substantiation selected by the applicant. When selecting the damage tolerance or safe life approach, attention should be given to geometry, inspectability, good design practice, and the type of damage/degradation of the structure under consideration.
6.2 Damage Tolerance (Fail-Safe) Evaluation
6.2.1 Structural details, elements, and subcomponents of critical structural areas should be tested under repeated loads to define the sensitivity of the structure to damage growth. This testing can form the basis for validating a no-growth approach to the damage tolerance requirements. The testing should assess the effect of the environment on the flaw growth characteristics and the no-growth validation. The environment used should be appropriate to the expected service usage. The repeated ] [ loading should be representative of anticipated service usage. The repeated load testing should include damage levels (including impact damage) typical of those that may occur during fabrication, assembly, and in service, consistent with the inspection techniques employed. The damage tolerance test articles should be fabricated and assembled in accordance with production specifications and processes so that the test articles are representative of production structure.
6.2.2 The extent of initially detectable damage should be established and be consistent with the inspection techniques employed during manufacture and in service. Flaw/damage growth data should be obtained by repeated load cycling of intrinsic flaws or mechanically introduced damage. The number of cycles applied to validate a no-growth concept should be statistically significant, and may be determined by load and/or life considerations. The growth or no growth evaluation should be performed by analysis supported by test evidence, or by tests at the coupon, element or subcomponent level.
6.2.3 The extent of damage for residual strength assessments should be established. Residual strength evaluation by component or subcomponent testing or by analysis supported by test evidence should be performed considering that damage. The evaluation should demonstrate that the residual strength of the structure is equal to or greater than the strength required for the specified design loads (considered as ultimate). It should be shown that stiffness properties have not changed beyond acceptable levels. For the no-growth concept, residual strength testing should be performed after repeated load cycling.
6.2.4 An inspection programme should be developed consisting of frequency, extent, and methods of inspection for inclusion in the maintenance plan. Inspection intervals should be established such that the damage will be detected between the time it initially becomes detectable and the time at which the extent of damage reaches the limits for required residual strength capability. For the case of no-growth design concept, inspection intervals should be established as part of the maintenance programme. In selecting such intervals the residual strength level associated with the assumed damage should be considered.
6.2.5 The structure should be able to withstand static loads (considered as ultimate loads) which are reasonably expected during the completion of the flight on which damage resulting from obvious discrete sources occur (i.e. uncontained engine failures, etc.). The extent of damage should be based on a rational assessment of service mission and potential damage relating to each discrete source.
6.2.6 The effects of temperature, humidity, and other environmental factors which may result in material property degradation should be addressed in the damage tolerance evaluation.
6.3 Fatigue (Safe-Life) Evaluation. Fatigue substantiation should be accomplished by component fatigue tests or by analysis supported by test evidence, accounting for the effects of the appropriate environment. The test articles should be fabricated and assembled in accordance with production specifications and processes so that the test articles are representative of production structure. Sufficient component, subcomponent, element or coupon tests should be performed to establish the fatigue scatter and the environmental effects. Component, subcomponent and/or element tests may be used to evaluate the fatigue response of structure with impact damage levels typical of those that may occur during fabrication, assembly, and in service, consistent with the inspection procedures employed. The component fatigue test may be performed with an as-manufactured test article if the effects of impact damage are reliably predicted by subcomponent and/or element tests and are accounted for in the fatigue test or in analysis of the results of the fatigue test. It should be demonstrated during the fatigue tests that the stiffness properties have not changed beyond acceptable levels. Replacement lives should be established based on the test results. An appropriate inspection programme should be provided.
7 Proof of Structure - Flutter. The effects of repeated loading and environmental exposure on stiffness, mass and damping properties should be considered in the verification of integrity against flutter and other aeroelastic mechanisms. These effects may be determined by analysis supported by test evidence, or by tests of the coupon, element or subcomponent level. ]
[ 8 Additional Considerations
8.1 Impact Dynamics. The present approach in airframe design is to assure that occupants have every reasonable chance of escaping serious injury under realistic and survivable impact conditions. Evaluation may be by test or by analysis supported by test evidence. Test evidence includes, but is not limited to, element or subcomponent tests and service experience. Analytical comparison to conventional structure may be used where shown to be applicable.
8.2 Flammability. (See appropriate JAR requirements in Table 1 of this ACJ.)
8.3 Lightning Protection. (See appropriate JAR requirements in Table 1 of this ACJ.)
8.4 Protection of Structure. Weathering, abrasion, erosion, ultraviolet radiation, and chemical environment (glycol, hydraulic fluid, fuel, cleaning agents, etc.) may cause deterioration in a composite structure. Suitable protection against and /or consideration of degradation in material properties should be provided for and demonstrated by test.
8.5 Quality Control. An overall plan should be established and should involve all relevant disciplines (i.e. engineering, manufacturing and quality control). This quality control plan should be responsive to special engineering requirements that arise in individual parts or areas as a result of potential failure modes, damage tolerance and flaw growth requirements, loadings, inspectability, and local sensitivities to manufacture and assembly.
8.6 Production Specifications. Specifications covering material, material processing, and fabrication procedures should be developed to ensure a basis for fabricating reproducible and reliable structure. The discrepancies permitted by the specifications should be substantiated by analysis supported by test evidence, or tests at the coupon, element or subcomponent level.
8.7 Inspection and Maintenance. Maintenance manuals developed by manufacturers should include appropriate inspection, maintenance and repair procedures for composite structures.
8.8 Substantiation of Repair. When repair procedures are provided in maintenance documentation, it should be demonstrated by analysis and/or test, that methods and techniques of repair will restore the structure to an airworthy condition. ]
TABLE 1
ACJ Paragraphs and related JAR texts

ACJ Paragraphs JAR-25 Paragraphs
1 Purpose No relevant JAR paragraph
2 Definitions No relevant JAR paragraph
3 General No relevant JAR paragraph
4 Material and Fabrication Development 25.603
25.605
[ 25.613
25.619 ]
5 Proof of Structure - Static 25.305
25.307(a)
6 Proof of Structure - Fatigue/Damage Tolerance 25.571
7 Proof of Structure - Flutter 25.629
8 Additional Considerations
8.1 Impact Dynamics 25.561
25.601
25.721
25.783(c) and (g)
25.785
25.787(a) and (b)
25.789
25.801
25.809
25.963(d) and (e)
8.2 Flammability 25.609(a)
25.853
25.855
25.859
25.863
25.865
25.867
25.903(c)(2)
25.967(e)
25.1121(c)
25.1181
25.1182
25.1183
25.1185
25.1189(a)(2)
25.1191
25.1193(c), (d) and (e)
8.3 Lightning Protection [ 25.581 (see ACJ 25X899 Paragraph 6) ]
25.609
[ 25X899 (see ACJ 25X899 Paragraph 6) ]
[ 25.954 (see ACJ 25X899 Paragraph 6) ]
8.4 Protection Structure 25.609
25.1529
8.5 Quality Control **
8.6 Production Specifications 25.603
25.605

ACJ 25.607(a)
Fasteners (Acceptable Means of Compliance)
See JAR 25.607(a)
In control systems where the means of connecting parts (e.g. a bolt) is assumed not to fail or become disconnected, it should be provided with secondary means of retention so that when the connection is installed the secondary means of retention becomes automatically effective in preventing the connection from moving out of position even though its primary means of retention (e.g. a nut) may have been omitted.
NOTE: A permanently locked (e.g. bench riveted with adequate dimensional control) connection, which is never broken down when assembled in the aeroplane, is not considered as a removable fastener.

ACJ 25.609
Protection of Structure (Acceptable Means of Compliance)
See JAR 25.609
The comprehensive and detailed national material standards accepted in the participating countries will be accepted as satisfying the requirement of JAR 25.609.

ACJ 25.631
Bird Strike Damage (Interpretative Material)
See JAR 25.631
Consideration should be given in the early stages of the design to the installation of items in essential services, such as control system components, and items which, if damaged, could cause a hazard, such as electrical equipment. As far as practicable, such items should not be installed immediately behind areas liable to be struck by birds.

ACJ 25.671(a)
Control Systems - General (Interpretative Material)
See JAR 25.671(a)
Control systems for essential services should be so designed that when a movement to one position has been selected, a different position can be selected without waiting for the completion of the initially selected movement, and the system should arrive at the finally selected position without further attention. The movements which follow and the time taken by the system to allow the required sequence of selection should not be such as to adversely affect the airworthiness of the aeroplane.

ACJ 25.671(b)
Control Systems - General (Interpretative Material)
See JAR 25.671(b)
For control systems which, if incorrectly assembled, would hazard the aeroplane, the design should be such that at all reasonably possible break-down points it is mechanically impossible to assemble elements of the system to give -
a. An out-of-phase action,
b. An assembly which would reverse the sense of the control, and
c. Interconnection of the controls between two systems where this is not intended.
Only in exceptional circumstances should distinctive marking of control systems be used to comply with the above.

ACJ 25.671(c)(1)
Control Systems - General (Interpretative Material)
See JAR 25.671(c)(1)
To comply with JAR 25.671(c)(1) there should normally be -
a. An alternative means of controlling the aeroplane in case of a single failure, or
b. An alternative load path.
However, where a single component is used on the basis that its failure is extremely improbable, it should comply with JAR 25.571(a) and (b).

ACJ 25.672(c)(1)
Stability Augmentation and Automatic and Power-operated Systems (Interpretative Material)
See JAR 25.672(c)(1)
The severity of the flying quality requirement should be related to the probability of the occurrence in a progressive manner such that probable occurrences have not more than minor effects and improbable occurrences have not more than major effects.
[ ACJ 25.679(a)(2)
Control System Gust Locks (Interpretative Material)
See JAR 25.679(a)(2)
If the device required by JAR 25.679(a) limits the operation of the aeroplane by restricting the movement of a control that must be set before take-off (e.g. throttle control levers), this device should be such that it will perform the function for which it is designed even when subject to likely maladjustment or wear, so that -
a. The movement of that control is restricted as long as the device is engaged; and
b. The movement of that control is unrestricted when the device is disengaged. ]
[ ACJ 25.679(b)
Control System Gust Locks (Interpretative Material)
See JAR 25.679(b)
For the purposes of meeting the design intent of this paragraph, flight means the time from the moment the aircraft first moves under its own power for the purpose of flight until the moment it comes to rest after landing. ]

ACJ 25.685(a)
Control System Details (Interpretative Material)
See JAR 25.685(a)
In assessing compliance with JAR 25.685(a) account should be taken of the jamming of control circuits by the accumulation of water in or on any part which is likely to freeze. Particular attention should be paid to the following:
a. The points where controls emerge from pressurised compartments.
b. Components in parts of the aeroplane which could be contaminated by the water systems of the aeroplane in normal or fault conditions; if necessary such components should be shielded.
c. Components in parts of the aeroplane where rain and/or condensed water vapour can drip or accumulate.
d. Components inside which water vapour can condense and water can accumulate.

ACJ 25.697(a)
Lift and Drag Devices, Controls (Acceptable Means of Compliance)
See JAR 25.697(a)
The lift-device operating control, in so far as selection of position is not automatic should be a lever working in a gated quadrant such that the pilot can readily and surely select by feel each take-off, en-route, approach and landing position established under JAR 25.101(d).
The lift-device operating control should be provided with a safety catch or similar device, operative at each position prescribed above, such that the action needed to release it is distinct from that used for releasing any other operating control operable by the same hand.
NOTE: Detents are not acceptable substitutes for gates at the intermediate positions for which gates are required.

ACJ 25.703(b)(4)
Take-off Warning System (Interpretative Material)
See JAR 25.703(b)(4)
1 Means provided to de-activate the warning should not be readily available to the flight crew during the take-off run.
2 Re-arming of the warning before each take-off may be accomplished either -
i. Automatically, or
ii. If the absence of re-arming is clear and unmistakable, manually.

ACJ 25.723(a)
Shock Absorption Tests (Interpretative Material)
See JAR 25.723(a)
[ 1 The prediction of landing loads should always be backed-up by energy absorption testing. However, it is acceptable to cover certain changes by calculations, providing the design concept is the same and the calculations have been shown to cover the test results realistically. For example, the following changes can be accepted in this sense: ]
[ a. Aeroplane sprung mass (effective weight) variations, including extrapolation from maximum landing weight to maximum take-off weight conditions.
b. Changes in shock absorber characteristics including pre-load, compression ratio and orifice sizes.
c. Changes in tyre characteristics.
d. Changes in unsprung mass (e.g. brakes).
2 The extent to which extrapolation can be accepted depends on the mathematical model employed and should therefore be subjected to negotiations with the Authority. ]

ACJ 25.729(e)
Retracting Mechanism (Interpretative Material)
See JAR 25.729(e)
[ 1 ] When light indicators are used, they should be arranged so that -
[ a. ] A green light for each unit is illuminated only when the unit is secured in the correct landing position.
[ b. ] A warning light consistent with JAR 25.1322 is illuminated at all times except when the landing gear and its doors are secured in the landing or retracted position.
[ 2 The warning required by JAR 25.729(e)(2) should preferably operate whatever the position of wing leading- or trailing-edge devices or the number of engines operating.
3 The design should be such that nuisance activation of the warning is minimised, for example -
a. When the landing gear is retracted after a take-off following an engine failure, or during a take-off when a common flap setting is used for take-off and landing;
b. When the throttles are closed in a normal descent; or
c. When flying at low altitude in clean or low speed configuration (special operation).
4 Inhibition of the warning above a safe altitude out of final approach phase either automatically or by some other means to prevent these situations is acceptable, but it should automatically reset for a further approach.
5 Means to de-activate the warning required by JAR 25.729(e) may be installed for use in abnormal or emergency conditions provided that it is not readily available to the flight crew, i.e. the control device is protected against inadvertent actuation by the flight crew and its de-activated state is obvious to the flight crew. ]

ACJ 25.729(f)
Protection of Equipment on Landing Gear and in Wheel Wells
(Acceptable Means of Compliance)
See JAR 25.729(f)
The use of fusible plugs in the wheels is not a complete safeguard against damage due to tyre explosion.
Where brake overheating could be damaging to the structure of, or equipment in, the wheel wells, an indication of brake temperature should be provided to warn the pilot.

ACJ 25.733(a)
Tyres (Interpretative Material)
See JAR 25.733(a)
The depth of tread below which friction characteristics are impaired should be specified and it should be possible to determine when the tread depth has worn below this limit.

ACJ 25.733(a)(1)
Tyres (Interpretative Material)
See JAR 25.733(a)(1)
The various materials should be so chosen that the tyres (whether re-treaded or not) are capable of the following:
a. Resisting deterioration, so far as practicable, resulting from the leakage of products used in the aeroplane such as hydraulic fluid, which might affect their serviceability.
b. Retaining sufficient strength and endurance when subjected to the environmental conditions existing in the landing-gear bay and to the heat generated by the normal use of the brakes, including the case of failure of the brake cooling system (if fitted).

ACJ 25.735(a)
Brakes (Interpretative Material)
See JAR 25.735(a)
The brakes and their various components should be appropriately protected against the ingress of such foreign bodies (water, mud, oil and other products) which may adversely affect their satisfactory performance. The shapes, sizes, and nature of the materials used in the manufacture of the various brake components should be so chosen that these materials will -
a. Withstand the hydraulic pressure and the loads to which they are subjected in normal conditions.
b. Withstand simultaneous application of normal and emergency pressures following a failure unless measures, agreed by the Certificating Authorities, have been taken to avoid such a contingency.
c. Maintain sufficient strength and endurance to withstand the temperature during normal braking of the aeroplane, even in the case of failure of the brake cooling system (if fitted).
d. Not induce, at any likely ground speed, vibrations likely to produce resonance in the aircraft structure or landing gear.

ACJ 25.735(b)
Brakes (Acceptable Means of Compliance and Interpretative Material)
See JAR 25.735(b)
The normal braking and emergency systems should be independent and supplied by separate power sources. After failure of the normal system, operation of the emergency system, and of the source of power supplying it, should be effected rapidly and safely either by the pilot or by an automatic device. In particular, all necessary steps should be taken to ensure that transfer from the normal system to the emergency system and generally from one braking system to another, while the brakes are not applied will not involve a risk of wheel locking. Unless equivalent safety is shown to be achieved by other means, power-operated installations should, either after the failure of any single source of hydraulic supply or of any single hydraulic component, or if no main or auxiliary power unit is operating, be capable of operation to the extent necessary for compliance with a. or b. as appropriate -
a. When no anti-skid device would be operating, provision for six applications of the brakes between fully off and the Normal Braking Force.
b. When an anti-skid device would be operating, provision for sufficient operation of the brakes to bring the aeroplane to rest when landing under the runway surface conditions for which the aeroplane is certificated.
NOTE: This recommendation is for the purpose of assessing the capacity of the emergency braking system.
Brake pipe protection. The run of the pipe-lines to the brakes should be such as to minimise the possibility of their being damaged by a burst tyre or shedding or flailing of tyre tread (see JAR 25.729(f)). The pipe lines should in any case be separated so that complete failure of the braking system would be improbable as a result of a single tyre failure.
Protection against fire. Unless it can be shown that hydraulic fluid which may be spilt on to hot brakes is unlikely to catch fire, the hydraulic system should be protected so as to limit the loss of fluid in the event of a serious leak. The precautions taken in the latter case should be such that the amount of fluid lost in the vicinity of the brakes is not sufficient to support a fire which is likely to hazard the aeroplane on the ground or in flight.

ACJ 25.735(c)
Brakes (Interpretative Material)
See JAR 25.735(c)
The braking force should increase or decrease progressively as the force or movement applied to the brake control is increased or decreased and the braking force should repsond to the control as quickly as is necessary for safe and satisfactory opertion. A brake control intended only for parking is excepted.

ACJ 25.735(d)
Brakes (Interpretative Material)
See JAR 25.735(d)
The control of the parking brake system should be protected against inadvertent operation during ground handling of the aeroplane and should be fitted with a suitable guard to prevent such operation. Alternatively, means should be provided to warn the pilot of possible braking during take-off. The parking brake control, whether or not it is independent of the emergency brake control, should be marked with the words 'Parking Brake' and should be constructed in such a way that, once operated, it remains locked in the 'on' position.

ACJ 25.735(e)
Brakes (Interpretative Material)
See JAR 25.735(e)
The anti-skid device should be designed to be no less reliable than the rest of the braking system. No single failure, occurring without the pilot's control being applied, in the hydraulic or electrical supply of the anti-skid valves should be the cause of wheel locking. In the event of any probable failure causing the poor operation of the anti-skids units, an automatic device, or a device which could be controlled by the pilot, should cut-out their action on the brakes and cut-in the appropriate circuit allowing the aeroplane to brake without making use of the anti-skid units. Should the supply for the anti-skid units be electrical, any failure in this supply should be indicated to the pilot.

ACJ 25.745(a)
Nose-wheel Steering (Interpretative Material)
See JAR 25X745(a)
In a powered nose-wheel steering system the normal supply for steering should continue without interruption in the event of failure of any one power-unit. With the remaining power-units operating at ground idling condition, the power supply should be adequate -
a. To complete an accelerate-stop manoeuvre following a power-unit failure which occurs during take-off, and
b. To complete a landing manoeuvre following a power-unit failure which occurs during take-off or at any later stage of flight.

ACJ 25.745(c)
Nose-wheel Steering (Interpretative Material)
See JAR 25X745(c)
1 No failure or disconnection need be assumed in respect of parts of proven integrity e.g. a simple jack or manual selector valve, but slow leakage from pipe joints and fracture of pipes should be considered as probable failures.
2 In assessing where the inadvertent application of steering torque as a result of a single failure would lead to danger, allowance may be made for the pilot's instinctive reaction to the effects of the fault. However, dependent on the urgency and rapidity of warning of the failure given to the pilot, allowance should be made for a reaction time before it is assumed that the pilot takes any corrective action.

ACJ 25.773(b)(1)(ii)
Pilot Compartment View (Acceptable Means of Compliance)
See JAR 25.773(b)(1)(ii)
For windshields protected by the application of electrical heat, a nominal heating capacity of 70 W/dm2 would be adequate.

ACJ 25.775(d)
The Integrity of Pressure Containing Transparencies (Acceptable Means of Compliance)
See JAR 25.775(d)
1 Mounting of window panels
1.1 The design of the mounting should be such that it will not transmit any unacceptable loads to the panel resulting from the items listed in 1.2, over the most adverse ranges of loading and climatic conditions.
1.2 Major items to be considered in designing the mounting for suitability over the ranges of loading and climatic conditions are -
a. Deflection of the panel and mounting under pressure;
b. Deflection of the mounting structure as a result of fuselage deflection;
c. Differential contraction and expansion between the panel and the mounting; and
d. Deflection of the panel resulting from temperature gradient across the thickness of the panel.
2 Strength Analysis and Testing
2.1 General. Panels of conventional design complying with the recommendations of this ACJ 25.775(d) will normally be acceptable if the conditions required to be fulfilled for the strength analysis and the tests of this paragraph 2 are satisfactorily completed. By conventional design is meant a laminated or double pane panel in which the main pressure bearing plies or panes, upon which the main and residual strength capability depends, are made from either thermally toughened soda lime glass or from acrylic material. It also assumes for the laminated panel that the plies are bonded together with an interlayer material not having inferior qualities to poly-vinyl butyral (PVB). Any other combination of plies or panes or materials should be discussed with the Authority as regards the applicability of the strength factors specified in this paragraph 2.
2.2 Reliability and Functioning Tests. It is recommended that cyclic tests on the window assemblies under closely representative operating conditions should be made to ensure an acceptable level of reliability in service of the transparencies and any associated heating systems. Such tests should also ensure that the panel installation is satisfactory when subjected to the most adverse combinations of panel heat, ambient temperature and pressure loading.
2.3 Fatigue. The substantiation of the fatigue integrity of the pressure cabin will be expected to include the window assemblies and their surrounding structure.
2.4 Static Strength
2.4.1 General
a. Compliance with the analysis of 2.4.2 and the tests of 2.4.3 as appropriate will be accepted as establishing that the strength of the weakest panel, in the most adverse operating conditions will be sufficient to ensure the safety objectives of JAR 25.775(d) are met.
b. The panel should be subjected to the most adverse likely combination of loads resulting from -
i The maximum normal working differential pressure, and
ii The external aerodynamic pressures on the panel.
c. Failure of the panel is considered to have occurred when the panel fails to hold pressure or there is a drastic increase in deflection of the panel. The failure of the outer thin facing glass in a laminated panel in the pilot's cabin is not considered to be a panel failure, unless this results in an unacceptable loss of vision, taking into account all other transparencies in the pilot's cabin.
2.4.2 Analysis
a. Laminated and Double Pane Glass Panels. The complete panel should be designed to a minimum ultimate factor of 4 on the most adverse pressure loading given by 2.4.1(b) without rupture. This analysis should be based on the appropriate strength of the glass as declared by the material manufacturer under loading conditions sustained for at least 30 minutes. The panel assembly should be assumed to be maintained at its normal working temperature as given by the panel heating system, if installed, ambient temperature on the outside and cabin temperature on the inside. The most adverse likely ambient temperature should be covered.
b. Laminated and Double Pane Acrylic Panels (Stretched and Unstretched).
i The maximum working stress level over the complete panel assembly should be shown by supporting evidence not to exceed a value consistent with the avoidance of fatigue and stress crazing, making due allowance for deterioration resulting from weathering, minor damage and scratching in service and use of cleaner fluids, etc. This analysis should be based on the appropriate strength of the acrylic as declared by the material manufacturer under sustained loading with the panel assembly maintained at its normal working temperature as given by the panel heating system, if installed, ambient temperature on the outside and cabin temperature on the inside. The most adverse likely ambient temperature should be covered.
ii In interpreting JAR 25.775(d) regarding structural integrity following any single failure in the installation or associated systems, the importance of avoiding overheat conditions for acrylic materials must be strongly emphasised, particularly for stretched acrylics in relation to the relaxation temperature for the material.
2.4.3 Strength Tests
a. Laminated and Double Pane Glass Panels. With any one main pressure bearing glass ply failed, it should be demonstrated that the panel can withstand an average value of twice the load of 2.4.1(b) long enough to give assurance that the pilot will have time to reduce the cabin pressure to a level at which the panel can withstand twice this reduced cabin pressure. The panel should then withstand this reduced factored pressure for a period sufficient to give assurance that the aeroplane can safely complete the flight.
i This test should be done so as to be as representative as possible of a sudden failure of the main pressure bearing glass ply in flight under once times the load of 2.4.1(b), with the load being gradually increased thereafter to the fully factored condition. Ambient air temperature should be reproduced on the outside of the panel and cabin temperature on the inside. The most adverse likely ambient temperature should be covered. The panel assembly should be heated in the same way as in the aeroplane at the beginning of the test.
ii Three specimens are required for this test.
b. Laminated and Double Pane Acrylic Panels (Stretched and Unstretched)
i Complete Panel. The analysis of the distribution of the working stress level over the panel given by 2.4.2(b) should be verified by test measurement.
ii Damaged Panel. With any one main pressure bearing acrylic ply failed for a laminated panel assembly, or with the normal load bearing pane failed for a double pane panel, it should be demonstrated that the panel can withstand an average value of four times the load of 2.4.1(b) long enough to give assurance that the pilot will have time to reduce the cabin pressure to a level at which the panel can withstand four times this reduced cabin pressure. The panel should then withstand this reduced factored pressure for a period sufficient to give assurance that the aeroplane can safely complete the flight.
- The failed acrylic ply may be represented either by a broken pattern to give an adequate representation of the worst which could occur, or, perhaps conservatively, by omission of the failed ply in the test panel. In either case the test should be done so as to be as representative as possible of a sudden failure of the particular acrylic ply in flight under once times the load of 2.4.1(b), with the load being gradually increased thereafter to the fully factored condition. Ambient air temperature should be reproduced on the outside of the panel and cabin temperature on the inside. The most adverse likely ambient temperature should be covered.
The panel should be heated in the same way as on the aeroplane at the beginning of the test.
- Three specimens are required for this test.
- In the case of a laminated panel, in confirmation of the inherent fail-safe character of the design, it should be further demonstrated by suitable tests that a crack in a particular ply cannot propagate through into an adjacent ply.
- Additional testing may be required if the design incorporates bolt-holes, cut-outs or other forms of stress raisers, particularly if non-stretched acrylic material is used.
3 Heating Systems. In considering any single failure in the installation or in associated systems, fire risks should be taken into account.

ACJ 25.777(a)
Cockpit Controls (Interpretative Material)
See JAR 25.777(a)
If the operating controls intended for operation by one pilot during take-off, accelerate stop, balked landing and landing are so arranged that the sequences of control actions during these manoeuvres necessitate the pilot having to change hands on the control column, these sequences should not involve such rapid hand changes that controllability will be prejudiced.

ACJ 25.777(e)
Cockpit Controls (Interpretative Material)
See JAR 25.777(e)
Except where a smaller distance is shown to be adequate, the distance aft of the landing gear control should not be less than 10 inches.

ACJ 25.777(g)
Cockpit Control Knob Shape (Interpretative Material)
See JAR 25.777(g)
Cockpit control knobs should conform to the general shapes (but not necessarily the exact sizes or specific proportions) in the following figure:


ACJ 25.785(c)
Seats and Safety Belts (Acceptable Means of Compliance)
See JAR 25.785(c)
1 Sharp edges or excrescences on the seats or parts of the passenger accommodation which might prove a source of danger not only to the occupants of the seats but particularly to the occupant seated to the rear should be avoided. All surfaces of passenger accommodation and those areas of the seat back lying within the arc of travel of the head of an occupant seated to the rear and restrained by a safety belt should be smooth and of large radius.
2 The radius of the arc of travel, representing the extremity of the occupant's head, should be taken as 710 mm (28 in). This allows for tall occupants and stretch in the safety belt. The centre of the radius of the arc of travel should be taken as 460 mm (18 in) forward and upward of the junction of the seat back and bottom at 35 to the latter (see Figure 1).

FIGURE 1
3 If the top of the seat back occurs within the arc of travel of the head, it should be padded to at least 25 mm (1 in) radius with at least 125 mm (05 in) of firm padding.
4 Any other substantially horizontal members occurring within the areas defined by paragraph 2 should either be padded as recommended in paragraph 3 or should be so arranged that the head will be deflected past them rather than strike them a direct blow. The tops of vertical members occurring within these areas should be so protected as to be at least as safe as horizontal members. No member should occur where it might be struck by the throat.
5 Where practicable, it is recommended that seat backs should be pivoted so as to move forward under emergency alighting acceleration loads so that the occupant of the seat behind only strikes a glancing blow on the seat back.

ACJ 25.785(g)
Seats, Berths, Safety Belts and Harnesses (Interpretative Material)
See JAR 25.785(g)
Where there is a risk that a safety belt or harness might, when not in use, foul the controls or impede the crew, suitable stowage should be provided, unless it can be shown that the risk can be avoided by the application of suitable crew drills.

ACJ 25.787(b)
[ Stowage Compartments (Acceptable Means of Compliance) ]
See JAR 25.787(b)
[ For stowage compartments in the passenger and crew compartments it must be shown by analysis and/or tests that under the load conditions as specified in JAR 25.561(b)(3), the retention items such as doors, swivels, latches etc., are still performing their retention function. In the analysis and/or tests the expected wear and deterioration should be taken into account. ]

ACJ 25.799
Water Systems (Acceptable Means of Compliance)
See JAR 25X799
Where water is provided in the aeroplane for consumption or use by the occupant, the associated system should be designed so as to ensure that no hazard to the aeroplane can result from water coming into contact with electrical or other systems.
Service connections (filling points) should be of a different type from those used for other services, such that water could not inadvertently be introduced into the systems for other services.

ACJ 25.807
Emergency Exit Access (Interpretative Material)
See JAR 25.807 and JAR 25.813
The term 'unobstructed' should be interpreted as referring to the space between the adjacent wall(s) and/or seat(s), the seatback(s) being in the most adverse position, in vertical projection from floor level to at least the prescribed minimum height of the exit.
[ ACJ 25.807(d) ]
Passenger Emergency Exits (Interpretative Material)
[seeJAR 25.807(d) ]
The optimum fore and aft location of Types I, II and III exits should be agreed between the applicant and the Certificating Authority bearing in mind the relevant considerations, including -
a. The varying likelihood of damage to different parts of the fuselage in emergency alighting conditions, and
b. The need to avoid the passengers having to evacuate the aeroplane where dangerous conditions (spilt fuel, hot engine parts, etc.) may exist.
[ ACJ 25.810(c)(2) ]
Emergency Evacuation (Acceptable Means of Compliance)
See JAR 25.810(c)(2)
Acceptable methods of measurement of reflectance are given in AC20-38A and AC20-47, published by the Federal Aviation Administration.
[ ACJ 25.811(e)(4)
Emergency Exit Marking (Interpretative Material)
See JAR 25.811(e)(4)
The indicating markings for all Type II and larger passenger emergency exit unlocking handle motions should conform to the general shapes and dimensions indicated by Figures 1 and 2.
NOTE: As far as is practicable the markings should be located to avoid obscuring viewing windows located on or alongside the exits, or coincidence with any other required marking or safety feature. ]
[ EXAMPLE MARKING FOR INDICATION OF LINEAR OPENING MOTION
Where practical and unambiguous arrow point and base of arrow shaft to be within 25 mm (1 inch) of fully unlocked and fully locked positions respectively

DIMENSIONS
A = 19 mm (075") minimum
B = 2 x A
C = B (recommended)
D = Indicative of the full extent of handle travel (each installation to be
individually assessed)
FIGURE 1
EXAMPLE MARKING FOR INDICATION OF ROTARY OPENING MOTION


Arrow point and base of arrow shaft to be within 25 mm (1 inch) of fully unlocked and fully locked positions respectively
DIMENSIONS
A = 19 mm (075") minimum
B = 2 x A
C = B (recommended)
D = Full extent of handle centreline travel
E = Three quarters of handle length (where practicable)
FIGURE 2 ]

ACJ 25.831(a)
Ventilation (Interpretative Material)
See JAR 25.831(a)
The supply of fresh air in the event of the loss of one source, should not be less than 04 lb/min per person for any period exceeding five minutes. However, reductions below this flow rate may be accepted provided that the compartment environment can be maintained at a level which is not hazardous to the occupant.

ACJ 25.831(c)
Ventilation (Interpretative Material)
See JAR 25.831(c)
1 To avoid contamination the fresh air supply should be suitably ducted where it passes through any compartment inaccessible in flight.
2 Where the air supply is supplemented by a recirculating system, it should be possible to stop the recirculating system and -
a. Still maintain the fresh air supply prescribed, and
b. Still achieve 1.

ACJ 25.851(a)
Fire Extinguishers (Interpretative Material)
See JAR 25.851(a)
1 Each extinguisher should be readily accessible and mounted so as to facilitate quick removal from its mounting bracket.
2 Unless an extinguisher is clearly visible, its location should be indicated by a placard or sign having letters of at least 0375 inches in height on a contrasting background. Appropriate symbols may be used to supplement such a placard or sign.
[ ACJ 25.851(a)(1) ]
Fire Extinguishers (Interpretative Material)
[ See JAR 25.851(a)(1) ]
1 The number and location of hand fire extinguishers should be such as to provide adequate availability for use, account being taken of the number and size of the passenger compartments and the location of toilets, galleys, etc. These considerations may result in the number being greater than the minimum prescribed.
2 Where only one hand extinguisher is required it should be located at the cabin attendant station, where provided, otherwise near the main entrance door.
3 Where two or more hand extinguishers are required and their location is not otherwise dictated by consideration of paragraph 1 above, an extinguisher should be located at each end of the cabin and the remainder distributed throughout the cabin as evenly as is practicable.
[ ACJ 25.851(a)(2) ]
Fire Extinguishers (Interpretative Material)
[ See JAR 25.851(a)(2) ]
There should be at least one fire extinguisher suitable for both flammable fluid and electrical equipment fires installed in each pilot's compartment. Additional extinguishers may be required for the protection of other compartments accessible to the crew in flight (e.g. electrical equipment bays) or from consideration of JAR 25.851(a)(2).
NOTE: Dry chemical fire extinguishers should not be used in pilot compartments because of the adverse effects on vision during discharge and, if non-conductive, interference with electrical contacts by the chemical residues.
[ ACJ 25.853(a)(1)
Flammability Standards for Materials Used in the Interiors of Aeroplanes with Passenger Capacity of 20 or More. (Acceptable Means of Compliance)
See JAR 25.853(a)(1)
Transparencies
JAR 25.853(a)(1) excludes lighting lenses from the requirements of Parts IV and V of Appendix F. In addition, windows and small transparent panels inserted within cabin partitions that afford flight attendants with a clear and unobstructed view of the cabin need not meet the requirements of this sub-paragraph, provided that they are of a small surface area. As guidance for the purpose of this ACJ, a small surface area is defined as being of less than 20 square feet (01858 sq. m.).
Galleys which are not isolated from the passenger cabin
Any galley surface that would be exposed to a cabin fire is considered an outer surface insofar as compliance with JAR 25.853(a)(1) is concerned.
Isolated compartments
Compartments which are isolated from the passenger cabin by a door or equivalent means need not meet the requirements of JAR 25.853(a)(1). ]

ACJ 25.857(b)
Cargo Compartment Classification (Acceptable Means of Compliance)
See JAR 25.857(b)
In showing compliance with JAR 25.857(b) consideration should be given to the effect of representative cargo loading conditions on the ability of the crew member to gain effective access to all parts of the compartment with a fire extinguisher. Where such access cannot be shown, it is recommended that means are provided to shut off all air supply within the compartment to increase the effectiveness of the extinguisher when used.

ACJ 25.857(d)
Cargo Compartment Classification (Interpretative Material)
See JAR 25.857(d)
1 The rate of ventilation and leakage into and out of Class D cargo compartments should be as low as practicable and should not exceed the value obtained from the following formula:
W = 2000 - V
where W = ventilation and leakage airflow in cu. ft./hour
V = compartment volume in cu. ft.
NOTE: Compliance with the ventilation rate should be shown by flight test in both pressurised and unpressurised flight. See also JAR 25.855(e).
See Orange Paper Amendment 96/1
2 When the volume of the compartment is greater than 1000 cu. ft. it should be demonstrated by a full scale test that a fire can safely be contained.
NOTE: The fire test may be waived if the cargo within the compartment is limited to container cargo and the containers are so constructed that they, in themselves, are individual Class D compartments.
3 The design of the compartment should be such that its mechanical integrity, including sealing, will be maintained when the compartment is exposed to those temperatures and pressures resulting from a fire within it and also in service, taking into account instructions contained in the service manual.
NOTE: The temperatures and pressures considered should, where appropriate, be those presented in FAA Reports RD 70-42, and 71-68 unless alternative values are substantiated by a fire test on the specific compartment design.

ACJ 25.863(a)
Flammable fluid fire protection (Interpretative Material)
See JAR 25.863(a)
The cooling air supply for any electrical or electronic equipment should be conveyed and discharged so as not to create a hazard following failure of the equipment.
NOTE: Where necessary the cooling duct should be fireproof.
Zones with surfaces which may be exposed to flammable fluids or vapours should be ventilated if the temperature of the surfaces may exceed (under normal or failure conditions) a dangerous value with regard to these fluids or vapours. Unless a higher value can be substantiated, a temperature exceeding 200C is considered dangerous.
[ ACJ 25.869 ]
Electrical System Fire and Smoke Protection (Interpretative Material and Acceptable Means of Compliance)
[ See JAR 25.869 ]
These requirements, and those of JAR 25.863 applicable to electrical equipment, may be satisfied by the following:
[ 1 ] Electrical components in regions immediately behind firewalls and in engine pod attachment structures should be of such materials and at such a distance from the firewall that they will not suffer damage that could hazard the aeroplane if the surface of the firewall adjacent to the fire is heated to 1100C for 15 minutes.
[ 2 ] Electrical equipment should be so constructed and/or installed that in the event of failure, no hazardous quantities of toxic or noxious (e.g. smoke) products will be distributed in the crew or passenger compartments.
[ 3 ] Electrical equipment, which may come into contact with flammable vapours should be so designed and installed as to minimise the risk of the vapours exploding under both normal and fault conditions. This can be satisfied by meeting the Explosion Proofness Standards of draft ISO document TC20/SC5/N.43, dated 1974.
[ ACJ 25.869(c) ]
Fire Protection for Oxygen Equipment (Interpretative Material)
[ See JAR 25.869(c) ]
1 High-pressure shut-off valves should be designed to provide effective slow opening and closing, so as to avoid the possible risk of fire or explosion.
2 Re-charging systems, if installed, should be provided with means to prevent excessive rates of charging which could result in dangerously high temperatures within the system. The charging system should also provide protection from contamination.
3 The compartments in which high-pressure system components, including source(s) are located should have adequate ventilation to ensure the rapid dilution of leaked oxygen. Such compartments should also provide adequate protection against contamination by liquids and other products which could result in the risk of fire.
4 Where in-situ charging facilities are provided, the compartments in which they are located should be accessible from outside the aircraft and as remote as possible from other service points and equipment. Placards should be provided, located adjacent to the servicing point, with adequate instructions covering the precautions to be observed when the system is being charged.
5 The installation of the system should be such that components and pipe lines -
a. Are adequately separated from electrical and fluid systems,
b. Are routed so as to minimise joints and sharp bends,
c. Are clear of moving controls and other mechanisms,
d. Are protected against grease or other lubricants, and are protected against the effects of vibration.
In addition, joints should where possible, be assembled dry, but where compounds are used for sealing they should be approved for that purpose.
6 Where the oxygen is supplied from chemical generators, the effects of heat emission, during both normal and inadvertent operation, on both the installation and other adjacent equipment, should be taken into account.

ACJ 25.899
Electrical Bonding and Protection against Lightning and Static Electricity (Interpretative Material and Acceptable Means of Compliance)
See JAR 25X899
1 Protection against Lightning Discharges. The aeroplane should be provided with means to conduct lightning strikes so that the aeroplane and its occupants will not be endangered. The means provided should be such as to -
a. Minimise damage to the aeroplane structure and components,
b. Prevent the dangerous malfunctioning of the aeroplane and its equipment as a result of the passage of lightning currents,
c. Prevent the occurrence of high potential differences within the aeroplane.
2 Characteristics of Lightning Discharges
2.1 The data contained in Table 1 should be used for the purpose of assessing the adequacy of lightning discharge protection of aeroplanes.
TABLE 1
Charge transfer maximum 600 coulombs
normal 50 to 200 coulombs
Peak current maximum 500 000 amperes
normal about 50 000 amperes
Duration of flash maximum 2 seconds
Duration of peak about 25 micro-seconds to half
current peak value critically damped
NOTE: The duration of flash may be made up of a number of discharges.
2.2 Where confirmatory tests are agreed with the Authority as being required to show compliance with the requirements of paragraph 1, then a discharge current having two components as given in Table 2, should be taken as being equivalent to a lightning strike from the aspects of heating and disruptive forces.
NOTE: This test is equally acceptable to that specified in FAA Advisory Circular AC 20-53 for fuel tank access doors and filler caps.
TABLE 2

Component Peak current Duration Charge transfer
____________________________________________________________________________________

1 200 000 amperes To crest value in 15 micro- 4 coulombs
seconds decaying to 50 000 amperes in 30 micro-seconds from initiation
_____________________________________________________________________________________

2 500 amperes 1 second rectangular wave 500 coulombs

NOTE: The above table gives typical test values, but the choice of either or both components for testing a given item of equipment will depend upon the relevance of these test components to the hazard in each particular case, such test conditions should be agreed with the Authority.
3 Protection against the Accumulation of Static Charges
3.1 General. All items, which by the accumulation and discharge of static charges may cause a danger of electrical shock, ignition of flammable vapours or interference with essential equipment (e.g. radio communications and navigational aids) should be adequately bonded to the main earth systems.
3.2 Intermittent Contact. The design should be such as to ensure that no fortuitous intermittent contact can occur between metallic and/or metallized parts.
3.3 High Pressure Refuelling and Fuel Transfer. Where provision is made for high pressure refuelling and/or for high rates of fuel transfer it should be established, by test, or by consultation with the appropriate fuel manufacturers, that dangerously high voltages will not be induced within the fuel system. If compliance with this requirement involves any restriction on the types of fuel to be used or on the use of additives, this should be established.
3.3.1 With standard refuelling equipment and standard aircraft turbine fuels, voltages high enough to cause sparking may be induced between the surface of the fuel and the metal parts of the tank at refuelling rates above approximately 250 gal/min. These induced voltages may be increased by the presence of additives and contaminants (e.g. anti-corrosion inhibitors, lubricating oil, free water), and by splashing or spraying of the fuel in the tank.
3.3.2 The static charge can be reduced as follows:
a. By means taken in the refuelling equipment such as increasing the diameter of refuelling lines and designing filters to give the minimum of electrostatic charging, or
b. By changing the electrical properties of the fuel by the use of anti-static additives and thus reducing the accumulation of static charge in the tank to negligible amount.
3.3.3 The critical refuelling rates are related to the aeroplane refuelling installations, and the designer should seek the advice of fuel suppliers on this problem.
4. Primary and Secondary Bonding Paths
4.1 Primary bonding paths are those paths which are required to carry lightning discharge currents. These paths should be of as low an electrical impedance as is practicable. Secondary bonding paths are those paths provided for other forms of bonding.
4.2 Where additional conductors are required to provide or supplement the inherent primary bonding paths provided by the structure or equipment, then the cross-sectional area of such primary conductors made from copper should be not less than 3 mm2 except that, where a single conductor is likely to carry the whole discharge from an isolated section, the cross-sectional area would be not less than 6 mm2. Aluminium primary conductors should have a cross-sectional area giving an equivalent surge carrying capacity.
4.3 Primary bonding paths should be used for -
a. Connecting together the main earths of separable major components which may carry lightning discharges,
b. Connecting engines to the main earth,
c. Connecting to the main earth all metal parts presenting a surface on or outside of the external surface of the aeroplane, and
d. Conductors on external non-metallic parts.
4.4 Where additional conductors are required to provide or supplement the inherent secondary bonding paths provided by the structure or equipment then the cross-sectional area of such secondary conductors made from copper should be not less than 1 mm2. Where a single wire is used its size should be not less than 12 mm diameter.
5 Resistance and Continuity Measurements. Measurements should be made to determine the efficacy of the bonding and connection between at least the following:
5.1 Primary Bonding Paths.
5.1.1 The extremities of the fixed portions of the aeroplane and such fixed external panels and components where the method of construction and/or assembly leads to doubt as to the repeatability of the bond, e.g. removable panels.
5.1.2 The engines and the main aeroplane earth.
5.1.3 External movable metal surfaces or components and the main aeroplane earth.
5.1.4 The bonding conductors of external non-metallic parts and the main aeroplane earth.
5.1.5 Internal components for which a primary bond is specified and the main aeroplane earth.
5.2 Secondary Bonding Paths.
5.2.1 Metallic parts, normally in contact with flammable fluids, and the main aeroplane earth.
5.2.2 Isolated conducting parts subject to appreciable electrostatic charging and the main aeroplane earth.
5.2.3 Electrical panels and other equipment accessible to the occupants of the aeroplane and the main aeroplane earth.
5.2.4 Earth connections, which normally carry the main electrical supply and the main aeroplane earth. The test on these connections should be such as to ensure that the connections can carry, without risk of fire or damage to the bond, or excessive volt drop, such continuous normal currents and intermittent fault currents as are applicable.
5.2.5 Electrical and electronic equipment and the aeroplane main earth, where applicable, and as specified by the aeroplane constructor.
5.2.6 Static discharger wicks and the main aeroplane structure.
[ 6 Electrical Properties of Composite Structure
6.1 In the case of lightning protection, for the partial conductors the method of surface protection will vary with the criticality of the structure in question. Deterioration of the means of protection or possible hidden damage to the material which may affect its structural integrity, need to be considered. While such materials provide a measure of electro-magnetic screening, the need for additional measures will be a function of the location of the material in relation to critical equipment and wiring in the aircraft. Particular attention will also have to be given to the protection required near fuel systems - e.g. fuel tanks.
For non-conducting materials which have no intrinsic lightning protection or screening properties, the measures taken will again depend on the relative locations of the material and critical systems or fuel and the possible loss of the components due to internal air pressures in the event of a strike.
6.2 The partial conducting materials should present no problem in dissipating P-static but problems can arise with the non-conductors. Depending upon the location of the material, protection may be required.
6.3 Electrical currents, other than lightning, can flow in some partial conducting materials and means may be required to limit this by provision of alternative current paths if the effect of large voltage drop is important or if such currents can damage the material.
6.4 Particular care has to be taken that all joints, permanent and temporary, are capable of carrying any currents which may flow particularly those resulting from lightning strikes. Structural damage and loss of screening capabilities may occur if these are not adequately controlled.
6.5 The adequacy of the material in supplying a ground plane for antenna may have to be considered. Again it will vary with the material and the radio frequency of the system. ]
ACJ - Subpart E

ACJ 25.901(b)(2)
Assembly of Components (Interpretative Material)
See JAR 25.901(b)(2)
The objectives of JAR 25.671(b) should be satisfied with respect to powerplant systems, where the safety of the aeroplane could otherwise be jeopardised.

ACJ 25.901(b)(4)
Electrical Bonding (Interpretative Material)
See JAR 25.901(b)(4)
Where the engine is not in direct electrical contact with its mounting, the engine should be electrically connected to the main earth system by at least two removable primary conductors, one on each side of the engine.

ACJ 25.901(e)
General (Interpretative Material)
See JAR 25.901(e)
See Orange Paper Amendment 96/1
The need for additional tests, if any, in hot climatic conditions should take account of any tests made by the engine constructor to establish engine performance and functioning characteristics and of satisfactory operating experience of similar power units installed in other types of aeroplane.
The applicant should declare the maximum climatic conditions for which compliance will be established and this should not be less severe than the ICAO Intercontinental Maximum Standard Climate (100F (378C) at sea level). If the tests are conducted under conditions which deviate from the maximum declared ambient temperature, the maximum temperature deviation should not normally exceed 25F (1388C).
As part of the tests referred to in JAR 25.901(e), the maximum drainage period following an abortive start before a further attempt to restart should be established. It should be demonstrated that successive attempts to restart do not create a fire hazard.
[ ACJ 25.903(a)
Engines (Acceptable Means of Compliance)
See JAR 25.903(a)
Acceptance of the original type certificate of the engine by other Airworthiness Authorities will depend upon the basis of its type certification as follows:
a. Engines Type Certificated to JAR-E. An engine type certificated to the applicable issue of JAR-E will be acceptable to other Participating Authorities in accordance with the provisions of the Arrangements Document.
b. Engines not Type Certificated to JAR-E. An engine type certificated to a code other than JAR-E will need to be found acceptable to each Airworthiness Authority in accordance with its national regulations. This may include showing compliance with the appropriate issue of JAR-E. ]
ACJ No. 1 to JAR 25.903 (d)(1)
Torching Flames (Acceptable Means of Compliance)
See JAR 25.903 (d)(1)
Where design precautions to minimise the hazard in the event of a combustion chamber burnthrough involve the use of torching flame resistant components and/or materials, satisfaction of the standards prescribed in British Standards Institution Specification 3G100: Part 2: Section 3: Sub-section 3.13, dated December 1973, is acceptable.
ACJ No. 2 to JAR 25.903 (d)(1)
Uncontained Engine Rotor Failures (Acceptable Means Of Compliance and Interpretative Material)
See JAR 25.903 (d)(1)
1 Turbine Engine Installations. Where containment of engine rotor debris has not been established, the following material provides a basis on which compliance may be shown with JAR 25.903 (d)(1).
2 Aeroplane Design Considerations
2.1 All practical design precautions should be taken to minimise, on the basis of good engineering judgement the risk of catastrophic damage due to non-contained engine rotor debris. This should include the position of the engine with respect to critical components or regions of the aeroplane such as -
a. The other engine(s) (especially those located on the same side of the aeroplane);
b. Fuselage pressurised hull and other primary structure;
c. Flight deck region;
d. Fuel system/tanks. (Consideration should be given to spillage of fuel into the engine compartment and any other region of the aeroplane where a fire hazard could result);
e. Essential control systems, including primary flight controls, electrical systems, hydraulic systems and shut-off means;
f. Engine fire extinguisher systems; and
g. Instrumentation essential for continued safe flight.
2.2 Practical design measures to minimise the risk of catastrophic damage may include for example, location of critical components or systems outside the vulnerable areas; duplication and adequate separation of critical components of systems, and/or protection by substantial airframe structure, taking account of the possible risk of simultaneous damage caused by the release (in random directions) of single fragments; location of shut-off means so that flammable fluids can be isolated in the event of damage to the system; use of protective armour or deflection shields; precautions to ensure that flammable fluids released from damaged lines or other components are not likely to contact possible ignition sources; possible redundant design or crack stoppers to limit the dynamic propagation of tears which have been caused by debris impact.
2.3 Where protection by substantial airframe structure or by protective armour or deflection shields is claimed, the adequacy of protection should be demonstrated by tests and/or analysis based on test data, using the criteria of the engine failure model of paragraph 3.
3 Engine Failure Model. The safety analysis required in paragraph 4 should be made using the following engine failure model unless, for the particular engine type concerned, evidence can be produced based on operating experience or engine design features to justify a different model.
3.1 Single One-Third Piece of Disc. It should be assumed that the one-third piece of disc has the maximum dimension corresponding to one-third of the disc with one-third blade height and an angular spread of 3 relative to the plane of rotation of the disc. Where energy considerations are relevant, the mass should be assumed to be one-third the bladed disc mass and its energy the translational energy (i.e. neglecting rotational energy) of the sector. (See Figure 1.)
3.2 Small Pieces of Debris. It should be assumed that the small piece of debris has a maximum dimension corresponding to one-third the bladed disc radius and an angular spread of 5 relative to the plane of the disc. Where energy considerations are relevant, the mass should be assumed to be th of the bladed disc mass and its energy the translational energy (neglecting rotational energy) of the piece travelling at rim speed (see Figure 2).
3.3 Alternative Engine Failure Model. For the purpose of the analysis, as an alternative to the engine failure model of paragraphs 3.1 and 3.2, the use of a single one-third piece of disc having an axial spread angle of 5 would be acceptable, provided that the objectives of paragraphs 2.1, 2.2 and 4.3 a. are satisfied.
4 Means of Compliance - Safety Analysis
4.1 An analysis should be made using the engine model defined in paragraph 3 to determine the critical areas of the aeroplane likely to be damaged by rotor debris and to evaluate the consequences. This should be determined in relation to the most critical flight phases.
4.1.1 A minimum delay of at least 15 seconds but in any event not more than 60 seconds should be assumed for the emergency engine shut down drill depending on the circumstances resulting from non-containment, taking into account the various phases of flight, and the fact that damage due to non-containment could result in a considerable increase in flight crew work load and delay in starting any of the emergency drills, for example, where there may be a multiplicity of warnings which require analysis of the situation by the flight crew to determine the cause.
4.1.2 Some degradation of the flight characteristics of the aeroplane or operation of a system may be permissible subject to the safe continuation of the flight. Account should be taken of the behaviour of the aeroplane under asymmetrical engine thrust or power conditions together with any possible damage to the flight control system, and of the predicted aeroplane recovery manoeuvre.
4.2 Drawings showing the trajectory paths of engine debris relative to critical areas should be provided. The analysis should include at least the following:
[ a. Damage to primary structure including the pressure cabin, engine mountings and control surfaces. From the results of this analysis, any structural damage resulting from uncontained rotor debris should be considered to be catastrophic unless the residual strength and flutter criteria of ACJ 25.571(a), sub-paragraph 2.7.2 can be met without failure of any part of the structure essential for successful completion of the flight. ]
b. Damage to any other engines (the consequences of subsequent non-containment of debris from the other engine(s), need not be considered).
c. Damage to services and equipment essential for safe flight (including indicating and monitoring systems), particularly control systems for flight, engine power, engine fuel supply and shut-off means and fire indication and extinguishing systems.
d. Pilot incapacitance.

Where R = the disc radius
b = blade length

The CG is taken to lie on the maximum dimension as shown

FIGURE 1 - SINGLE ONE-THIRD DISC FRAGMENT

Where R = disc radius
b = blade length

Maximum dimension = (R + b)
Mass assumed to be th of bladed disc
CG is taken to lie on the disc rim
FIGURE 2 - SMALL PIECE OF DEBRIS
e. Penetration of the fuel system, where this could result in the release of fuel into personnel compartments or an engine compartment or other regions of the aeroplane where this could lead to a fire (or explosion).
f. Damage to the fuel system, especially tanks, resulting in the release of a large quantity of fuel.
g. Penetration and distortion of firewalls and cowling permitting a spread of fire.
NOTE: Consideration of the effect of damage should include degradation of the performance and handling characteristics of the aeroplane.
4.3 When all practical design precautions have been taken (see paragraph 2.1) and the safety analysis made using the engine failure model defined in paragraph 3 shows that catastrophic risk still exists for some components or systems of the aeroplane, the level of catastrophic risk should be evaluated. It is [ considered that the objective of the requirement will have been met if the levels of risk stated in a., b. and c., ] as appropriate, have been achieved.
NOTE: It is accepted that due allowance should be made for the size and broad configuration of the aeroplane and that this may prevent the prescribed levels of risk being achieved.
a. Single One-third Piece of Disc. There is not more than a 1 in 20 chance of catastrophe resulting from the release of a single one-third piece of disc as defined in paragraph 3.1.
b. Small Piece of Debris. There is not more than a 1 in 40 chance of catastrophe resulting from the release of a piece of debris as defined in paragraph 3.2.
[ c. Multiple Fragments. (Only applicable to any duplicated or multiplicated system where all of the system channels contributing to its function have some part which is within a distance equal to the diameter of the largest bladed rotor, measured from the engine centreline). There is not more than a 1 in 10 chance of catastrophe resulting from the release in three random directions of three one-third fragments of a disc each having a uniform probability of ejection over the 360 (assuming an angular spread of 3 relative to the plane of the disc) causing coincidental damage to systems which are duplicated or multiplicated.
NOTE: Where dissimilar systems can be used to carry out the same function (e.g. elevator control and pitch trim), they should be regarded as duplicated (or multiplicated) systems for the purpose of this sub-paragraph. ]
4.4 The aeroplane risk levels specified, resulting from the release of rotor debris, are the mean values obtained by averaging those for all discs on all engines of the aeroplane, assuming a typical flight. Individual discs or engines need not meet these risk levels nor need these risk levels be met for each phase of flight if either -
a. No disc shows a higher level of risk averaged throughout the flight greater than twice those stated in paragraph 4.3.
NOTE: The purpose of this paragraph is to ensure that a fault which results in repeated failures of any particular disc design, would have only a limited effect on aeroplane safety.
b. Where failures would be catastrophic in particular phases of flight only, allowance is made for this on the basis of conservative assumptions as to the proportion of failures likely to occur in these phases. A greater level of risk could be accepted if the exposure exists only during a particular phase of flight e.g. during take-off. The proportional risk of engine failure during the particular phases of flight is given in SAE Paper AIR 1537 dated October 1977 'Report on Aircraft Engine Containment'. See also data contained in the CAA paper 'Engine Non-Containment - The CAA View', which includes Figure 3. This paper is published in NASA Report CP-2017, 'An assessment of Technology for Turbo-jet Engine Rotor Failures', dated August 1977.

FIGURE 3 -ALL NON-CONTAINMENTS BY PHASE OF FLIGHT

ACJ 25.903(e)(2)
Engines (Interpretative Material)
See JAR 25.903(e)(2)
1 General
1.1 In general the relight envelope required in JAR 25.903(e)(2) may consist of two zones -
a. One zone where the engine is rotated by windmilling at or beyond the minimum rpm to effect a satisfactory relight, and
b. Another zone where the engine is rotated with assistance of the starter at or beyond the minimum rpm to effect a satisfactory relight.
1.2 The minimum acceptable relight envelope is defined in paragraph 2.
2 Envelope of Altitude and Airspeed
2.1 Sufficient flight tests should be made over the range of conditions detailed in 2.2 and 2.3, to establish the envelope of altitude and airspeed for reliable engine restarts, taking into account the results of restart tests completed by the engine constructor on the same type of engine in an altitude test facility or flying test bed, if available, and the experience accumulated in other aircraft with the same engine. The effect of engine deterioration in service should be taken into account.
2.2 Altitude and Configuration. From sea-level to the maximum declared restarting altitude in all appropriate configurations likely to affect restarting, including the emergency descent configuration.
2.3 Airspeed. From the minimum to the maximum declared airspeed at all altitudes up to the maximum declared engine restarting altitude. The airspeed range of the declared relight envelope should cover at least 30 kt.
2.4 Delay Tests. The tests referred to in paragraph 2.2 should include the effect on engine restarting performance of delay periods between engine shut-down and restarting of -
a. Up to two minutes, and
b. At least fifteen minutes or until the engine oil temperatures are stabilised at their cold soak value.

ACJ 25.905(a)
Propellers (Acceptable Means of Compliance)
See JAR 25.905(a)
Acceptance of the original type certificate of the propeller by other Airworthiness Authorities will depend upon the basis of its type certification as follows:
a. Propellers Type Certificated or Otherwise Approved to JAR-P. A propeller type certificated or otherwise approved to the applicable issue of JAR-P will be acceptable to other Participating Authorities in accordance with the provisions of the Arrangements Document.
b. Propellers not Type Certificated or Otherwise Approved to JAR-P. A propeller type certificated or otherwise approved to a code other than JAR-P will need to be found acceptable by each Airworthiness Authority in accordance with its national regulations. This may include showing compliance with the appropriate issue of JAR-P.

ACJ 25.905(d)
Release of Propeller Debris (Acceptable Means of Compliance and Interpretative Material)
See JAR 25.905(d)
1 Propeller Installation. Design features of the propeller installation, including its control system, which are considered to influence the occurrence of propeller debris release and/or mode of such a failure should be taken into account when assessing the aeroplane against JAR 25.905(d).
2 Aeroplane Design Conditions
2.1 Impact Damage Zone. All practical precautions should be taken in the aeroplane design to minimise, on the basis of good engineering judgement, the risk of Catastrophic Effects due to the release of part of, or a complete propeller blade. These precautions should be taken within an impact zone defined by the region between the surfaces generated by lines passing through the centre of the propeller hub making angles of at least five degrees forward and aft of the plane of rotation of each propeller. Within this zone at least the following should be considered.
a. The vulnerability of critical components and systems (e.g. location, duplication, separation, protection); and
b. The fire risk in the event of flammable fluid release in association with potential ignition sources (e.g. location, protection, shut-off means).
2.2 Other Considerations. Consideration should be given to the effects on the aeroplane resulting from -
a. The likely out of balance forces due to the release of part of, or a complete propeller blade; and
b. Loss of a complete propeller.

ACJ 25.929(a)
Propeller De-icing (Acceptable Means of Compliance)
See JAR 25.929(a)
[ Where the propeller has been fitted to the engine in complying with the tests of ACJ E 780, compliance with ] JAR 25.929(a) will be assured.

ACJ 25.939(a)
Turbine Engine Operating Characteristics (Interpretative Material)
See JAR 25.939(a)
The wording 'in flight' should be interpreted to cover all operating conditions from engine start until shut-down.
2 If the airflow conditions at the engine air intake can be affected by the operating conditions of an adjacent engine, the investigation should include an exploration of the effects of running the adjacent engine at the same and at different conditions over the whole range of engine operating conditions, including reverse thrust. An investigation of the effect of malfunctioning of an adjacent engine should also be included.

ACJ 25.939(c)
Turbine Engine Operating Characteristics (Acceptable Means of Compliance and Interpretative Material)
See JAR 25.939(c)
1 The investigation should cover the complete range, for which certification is required, of aeroplane speeds, attitudes, altitudes and engine operating conditions including reverse thrust, and of steady and transient conditions on the ground and in flight, including crosswinds, rotation, yaw and stall. Non-critical conditions of operation which need not be considered should be agreed with the Authority.
2 If the airflow conditions at the engine air intake can be affected by the operating conditions of an adjacent engine, the investigation should include an exploration of the effects of running the adjacent engine at the same and at different conditions over the whole range of engine operating conditions, including reverse thrust. An investigation of the effect of malfunctioning of an adjacent engine should also be included.
3 Compliance with the requirement may include any suitable one or combination of the following methods; as agreed with the Authority.
a. Demonstration that the variations in engine inlet airflow distortion over the range defined in 1 are within the limits established for the particular engine type.
b. An investigation of blade vibration characteristics by the method and of the scope indicated in [ JAR-E 650(d) and ACJ E 650(d) (except that Maximum Take-off rpm need not be exceeded) carried out on - ]
i A representative installation on the ground using test equipment where the actual conditions of operation in the aeroplane are reproduced, or
ii A representative aeroplane on the ground and in flight as appropriate to the conditions being investigated.
c. The completion of sufficient flying with representative installations prior to certification such as to demonstrate that the vibration levels are satisfactory.
d. Any other method acceptable to the Authority.

ACJ 25.939(d)
Turbine Engine Operating Characteristics (Acceptable Means of Compliance)
See JAR 25.939(d)
Compliance with JAR 25.939(d) may consist of flight tests using vibration measuring equipment on which engine test bed vibration levels were established, or the equipment intended to be supplied on production engines provided the Authority considers the equipment sensitive enough for the purpose of showing compliance with the requirements.

ACJ 25.954
Fuel System Lightning Protection (Acceptable Means of Compliance and Interpretative Material)
See JAR 25.954
1 The fuel storage system and the outlets of the venting and jettisoning systems of the aeroplane, should be so situated and/or protected, that the probability of a catastrophe being caused by them being struck by lightning is extremely improbable.
NOTE: The location of the fuel tanks and vents within the airframe may be such as to satisfy this.
2 In addition, the outlets of venting and jettisoning systems should be so located and designed that-
a. They will not, under any atmospheric conditions which the aeroplane may encounter, experience electrical discharges of such magnitudes as will ignite any fuel/air mixture of the ratios likely to be present, and
b. The fuel and its vapours in flammable concentrations will not pass close to parts of the aeroplane which will produce electrical discharges capable of igniting fuel/air mixtures.
NOTE: Electrical discharges may, in addition to direct lightning strikes, be caused by corona and streamer formation in the vicinity of thunderstorms.
3 The fuel system of the aeroplane should be so designed that the passage of lightning discharges through the main aeroplane structure will not produce, by the process of conduction or induction, such potential differences as will cause electrical sparking through areas where there may be flammable vapours.
NOTE: For aeroplanes of conventional shape, an acceptable method of complying with JAR 25.954 is given in FAA Advisory Circular AC20-53 - 'Protection of Aircraft Fuel Systems against Lightning'. For aeroplanes of non-conventional shape, re-definition of the zones may be necessary.
[ ACJ 25.955(a)(4)
Fuel Flow (Interpretative Material)
See JAR 25.955(a)(4)
The word "blocked" should be interpreted to mean "with the moving parts fixed in the position for maximum pressure drop". ]

ACJ 25.961(a)(5)
Fuel System Hot Weather Operation (Acceptable Means of Compliance)
See JAR 25.961(a)(5)
Subject to agreement with the Authority, fuel with a higher vapour pressure may be used at a correspondingly lower fuel temperature provided the test conditions closely simulate flight conditions corresponding to an initial fuel temperature of 110F at sea level.

ACJ 25.963(a)
Fuel Tanks: General (Interpretative Material)
See JAR 25.963(a)
Precautions should be taken against the possibility of corrosion resulting from microbiological contamination of fuel.

ACJ 25.963(d)
Fuel Tanks: General (Acceptable Means of Compliance)
See JAR 25.963(d)
Fuel tank installations should be such that the tanks will not be ruptured by the aeroplane sliding with its landing gear retracted, nor by a landing gear, nor an engine mounting tearing away.
Fuel tanks inboard of the landing gear or inboard of or adjacent to the most outboard engine, should have the [ strength to withstand fuel inertia loads appropriate to the accelerations specified in JAR 25.561(b)(3) considering the maximum likely volume of fuel in the tank(s). For the purposes of this substantiation it will not be necessary to consider a fuel volume beyond 85% of the maximum permissible volume in each tank. For calculation of inertia pressures a typical density of the appropriate fuel may be used. ]
[ ACJ 25.963(g)
Fuel Tanks: General (Acceptable Means of Compliance)
See JAR 25.963(g)
1 Purpose. This ACJ sets forth an acceptable means of showing compliance with the provisions of JAR-25 dealing with the certification requirements for fuel tank access covers. Guidance information is provided for showing compliance with the impact resistance requirements of 25.963(g).
2 Background. Fuel tank access covers have failed in service due to impact with high speed objects such as failed tyre tread material and engine debris following engine failures. Failure of an access cover on a wing fuel tank may result in the loss of hazardous quantities of fuel which could subsequently ignite.
3 Impact Resistance
a. All fuel tank access covers must be designed to minimise penetration and deformation by tyre fragments, low energy engine debris, or other likely debris, unless the covers are located in an area where service experience or analysis indicates a strike is not likely. The rule does not specify rigid standards for impact resistance because of the wide range of likely debris which could impact the covers. The applicant should however, choose to "minimise penetration and deformation" by testing covers using debris of a type, size, trajectory, and velocity that represents conditions anticipated in actual service for the aeroplane model involved. There should be no hazardous quantity of fuel leakage after impact. The access covers, however, need not be more impact resistant than the contiguous tank structure.
b. In the absence of a more rational method, the following criteria should be used for evaluating access covers for impact resistance.
i. Covers located within 15 inboard and outboard of the tyre plane of rotation, measured from the centre plane of tyre rotation with olco strut in the nominal position, should be evaluated. The evaluation should be based on the results of impact tests using tyre tread segments having width and length equal to the full width of the tread, with thickness of the full tread plus casing. The velocities used in the assessment should be based on the highest speed that the aircraft is likely to use on the ground. Generally, this will be the higher of the aircraft rotation speed (VR) and the flapless landing speed.
ii. Covers located within 15 forward of the front compressor or fan plane measured from the centre of rotation to 15 aft of the rearmost turbine plane measured from the centre of rotation, should be evaluated for impact from small fragments (shrapnel). The covers need not be designed to withstand impact from high energy engine fragments such as rotor segments. ]

ACJ 25.965(a)
Fuel Tank Tests (Interpretative Material)
See JAR 25.965(a)
The analysis or tests should be performed on each complete tank in the configuration ready and capable of flight. Each complete tank means any tank fully equipped which is isolated from other tanks by tank walls or which may be isolated by valves under some flight configurations.

ACJ 25.967(a)(3)
Fuel Tank Installation (Interpretative Material)
See JAR 25.967(a)(3)
The installation of a flexible tank and its venting, according to JAR 25.975(a)(3) should be such that the tank liner will not be deformed in such a way as to significantly affect the fuel quantity indication.

ACJ 25.979(d)
Pressure Fuelling Systems (Acceptable Means of Compliance)
See JAR 25.979(d)
1 Pressure fuelling systems, fuel tanks and the means preventing excessive fuel pressures, should be designed to withstand normal maximum fuelling pressure of not less than 345 kN/m2 (50 psi) at the coupling to the aeroplane.
2 Pressure fuelling systems should be so arranged that the fuel entry point is at or near the bottom of the tank so as to reduce the level of electrostatic charge in the tank during fuelling.

ACJ 25.1027
Inadvertent Propeller Feathering (Interpretative Material)
See JAR 25.1027
The design of the propeller feathering system should be such that it is possible to complete the feathering and the unfeathering operation under all normal operating conditions.

ACJ 25.1027(b)
Propeller Feathering (Interpretative Material)
See JAR 25.1027(b)
The amount of trapped oil should be sufficient to cover one feathering operation; taking into account the maximum oil leakage in the feathering system due to wear and deterioration in service.

ACJ 25.1041
Tests in hot climatic conditions (Interpretative Material)
See JAR 25.1041
See Orange Paper Amendment 96/1

ACJ 25.1091(d)(2)
Precipitation Covered Runways (Acceptable Means of Compliance)
See JAR 25.1091(d)(2)
1 Except where it is obvious by inspection or other means, that precipitation on the runway would not enter the engine air intake under the declared operating conditions, including the use of the thrust reverser, compliance with the requirements should be demonstrated by tests using tyres representative of those to be approved for operational use. These tests should clear the aeroplane for operation from runways which are normally clear and also for operation in precipitation up to 13 mm (05 in) depth of water or dense slush. The tests should be conducted with the minimum depth of 13 mm (05 in) and an average depth of 19 mm (075 in), [ or if approval is sought for a greater depth than 13 mm (05 in), the average depth should be 15 times the ] depth for which the take-offs are to be permitted, and the minimum depth should be not less than the depth for which take-offs are to be permitted.
2 It should be shown that the engines operate satisfactorily without unacceptable loss of power at all speeds from zero up to lift-off speed and in the attitudes likely to be used. Any special aeroplane handling techniques necessary to ensure compliance with the requirement should comply with the handling techniques assumed in establishing the scheduled performance of the aircraft.
3 The tests may be made in water or slush either be complete take-offs and landings as necessary in the specified precipitation conditions, or by a series of demonstrations in areas of precipitation sufficiently large to permit the spray pattern to become stabilised and to determine engine behaviour and response. Experience has shown that where a trough is used, a length of 70 to 90 m (230 to 295 ft) is usually satisfactory. If marginal results are obtains the effect of the difference between water and slush should be examined.
4 The effects of cross-winds should be examined and where necessary a cross-wind limitation established for inclusion in the Flight Manual for operation from precipitation covered runways.
5 It may be difficult to deduce the effect of low density precipitation (dry snow) from high density testing, but nevertheless clearance of the aeroplane for operation in dense precipitation up to 13 mm (05 in) will usually clear the aeroplane for operation in low density precipitation of depths greater than 10 cm (4 in) depth. If clearance is requested for operation in low density precipitation of depths greater than 10 cm (4 in) additional tests (in low density precipitation having a depth close to that for which approval is sought) will be necessary.
6 When auxiliary devices are fitted to prevent spray from being ingested by the engines it will be necessary to do additional tests in low density precipitation to permit operations in depths greater than 25 mm (1 in).

ACJ 25.1091(e)
Air Intake System (Interpretative Material)
See JAR 25.1091(e)
The parts or components to be considered are, for example, intake splitters, acoustic lining if in a vulnerable location and inlet duct-mounted instrumentation.

ACJ 25.1093(b)
Propulsion Engine Air Intakes (Acceptable Means of Compliance and Interpretative Material)
See JAR 25.1093(b)
1 General. Two ways of showing compliance with JAR 25.1093(b) are given.
1.1 Method 1. Method 1 is an arbitrary empirical method based on United Kingdom and French practice. This method is acceptable to all participating countries.
1.2 Method 2. Method 2 is a general approach based on US practice in applying FAR Part 25, Appendix C. If this method is used, each application will have to be evaluated on its merits.
2 Method 1 (Acceptable Means of Compliance)
2.1 In establishing compliance with the requirements of JAR 25.1093(b), reference should be made to ACJ 25.1419, paragraph 1.
2.2 The intake may be tested with the engine and propeller where appropriate in accordance with the [ requirements of JAR-E 780 and ACJ E 780. ]
2.3 When the intake is assessed separately (e.g. lack of suitable test facilities, change in the design of the intake, intake different from one tested with the engine) it should be shown that the effects of intake icing would not invalidate the engine tests of JAR-E. Factors to be considered in such evaluation are:
a. Distortion of the airflow and partial blockage of the intake.
b. The shedding into the engine of intake ice of a size greater than the engine is known to be able to ingest.
c. The icing of any engine sensing devices, other subsidiary intakes or equipment contained within the intake.
d. The time required to bring the protective system into full operation.
2.4 Tests in Ice-forming Conditions. An acceptable method of showing compliance with the requirements of JAR 25.1093(b), including Appendix C, is given in this paragraph.
2.4.1 When the tests are conducted in non-altitude conditions, the system power supply and the external aerodynamic and atmospheric conditions should be so modified as to represent the required altitude condition as closely as possible. The altitudes to be represented should be as indicated in Table 1 for simulated tests, or that appropriate to the desired temperature in flight tests, except that the test altitude need not exceed any limitations proposed for approval. The appropriate intake incidences or the most critical incidence, should be simulated.
2.4.2 A separate test should be conducted at each temperature condition of Table 1, the test being made up of repetitions of either the cycle -
a. 28 km in the conditions of Table 1 column (a) appropriate to the temperature, followed by 5 km in the conditions of Table 1 column (b) appropriate to the temperature, for a duration of 30 minutes, or
b. 6 km in the conditions of Table 1 column (a) appropriate to the temperature, followed by 5 km in the conditions of Table 1 column (b) appropriate to the temperature, for a duration of 10 minutes.
TABLE 1
Ambient air Altitude Liquid water content (g/m3) Mean effective
temperature droplet diameter (C) (ft) (m) (a) (b) (um)

-10 17000 5200 0.6 2.2

-20 20000 6100 0.3 1.7 20

-30 25000 7600 0.2 1.0
2.4.3 Either by separate tests, or in combination with those of 2.4.2 it should be demonstrated that the ice accretion is acceptable after a representative delay in the selection of the ice-protection systems, such as might occur during inadvertent entry into the conditions. In lack of other evidence a delay of two minutes (to switch on the system) should normally be achieved. The time for the system to warm up should be represented.
2.4.4 For each test, the ice protection supply should be representative of the minimum engine power for which satisfactory operation in icing conditions is claimed.
2.4.5 If at the conclusion of each of the tests of 2.4.2 there is excessive ice accretion then the heat flow and airflow should be changed simultaneously to simulate an engine acceleration to demonstrate the pattern of ice shedding, which should be acceptable to the engine.
2.4.6 Where the minimum engine power necessary to provide adequate protection (as established in 2.4.2) is greater than that required for descent, an additional test representative of the minimum engine power associated with descent should be conducted by means of either -
a. A run at the -10C condition of Table 1, column (a), for sufficient duration to cover an anticipated descent of 10 000 ft, or
b. A run simulating an actual descent, at the conditions of Table 1 column (a), covering an altitude change of not less than 10 000 ft, the highest total temperature reached being not more than 0C.
2.4.7 If at the conclusion of the test in 2.4.6 there is excessive ice accretion then the heat flow and airflow should be changed simultaneously to simulate an engine acceleration and the ambient temperature should be increased to above 0C to demonstrate the pattern of total ice shedding which should be acceptable to the engine.
2.4.8 If the intake contains features or devices which could be affected by freezing fog conditions then in addition to the above tests of 2.4.2, 2.4.3 and 2.4.6 a separate test on these parts should be conducted for a duration of 30 minutes, in an atmosphere of -2C and a liquid water concentration of 03g/m3, with the heat supply to the tested part as would be available with the engine set to the minimum ground idle conditions approved for use in icing. The mean effective droplet size for the test should be 20 m. At the end of the period the ice accretion on the tested part should not prevent its proper functioning, nor should the ice be of such size as to hazard the engine if shed.
3 Method 2 (Interpretative Material)
3.1 In establishing compliance with the requirements of JAR 25.1093(b), reference should be made to JAR 25.1419 and ACJ 25.1419.
3.2 The intake may be tested with the engine and propeller where appropriate in accordance with a programme of tests which results from an analysis of the icing conditions and the engine conditions appropriate to the installation.
3.3 When the intake is assessed separately it should be shown that the effects of intake icing would not invalidate any engine certification tests. Factors to be considered in such evaluation are -
a. Distortion of the airflow and partial blockage of the intake.
b. The shedding into the engine of intake ice of a size greater than the engine is known to be able to ingest.
c. The icing of any engine sensing devices, other subsidiary intakes or equipment contained within the intake.
d. The time required to bring the protective system into full operation.
3.4 When tests are conducted in non-altitude conditions, the system power supply and the external aerodynamic and atmospheric conditions should be so modified as to represent the altitude condition as closely as possible. The appropriate intake incidences or the most critical incidence, should be simulated.
3.5 Following the analysis required in JAR 25.1419(b), which will determine the critical icing conditions within the envelope of icing conditions defined by Appendix C Figures 1 to 3 and Appendix C Figures 4 to 6, tests should be conducted at such conditions as are required to demonstrate the adequacy of the design points.
3.6 It should be demonstrated that the ice accretion is acceptable after a representative delay in the selection of the ice protection systems, such as might occur during inadvertent entry into the conditions. In lack of other evidence a delay of two minutes (to switch on the system) should normally be achieved in continuous maximum icing conditions. The time for the system to warm up should be represented.
3.7 If at the conclusion of each of the tests there is excessive ice accretion then the heat flow and airflow should be changed simultaneously to simulate an engine acceleration to demonstrate the pattern of ice shedding, which should be acceptable to the engine.
3.8 Where the minimum engine power necessary for adequate protection as established above is greater than that required for descent, this should be considered in the analysis, and test evidence may have to be provided to demonstrate acceptability. The icing conditions and vertical extent are as in Figure 1 of Appendix C. Any ice able to be shed from the intake into the engine should be acceptable to the engine.
3.9 If the intake contains features or devices which could be affected by freezing fog conditions then a separate assessment for these parts should be conducted assuming a duration of 30 minutes and an atmosphere of -2C, and a liquid water concentration of 03g/m3, with the heat supply to the tested part as would be available with the engine set to the minimum ground idle conditions approved for use in icing. The mean effective droplet size should be 20m. At the end of the period the ice accretion on the part should not prevent its proper functioning, nor should the ice be of such size as to hazard the engine if shed.

ACJ 25.1103(d)
Air Intake System Ducts (Interpretative Material)
See JAR 25.1103(d)
For a single failure case leading to a fire and air duct rupture, consideration should be given to the possibility of fire aggravation due to air flowing into a designated fire zone of an engine from the remaining engine(s), or another source outside the affected fire zone.

ACJ 25.1121(a)
General (Interpretative Material)
See JAR 25.1121(a)
1 If necessary, each exhaust system should be provided with drains to prevent hazardous accumulation of fuel under all conditions of operation.
2 Tests should be made to demonstrate compliance with JAR 25.1121(a) and these should include engine starting in downwind conditions and thrust reversal.

ACJ 25.1121(b)
General (Interpretative Material)
See JAR 25.1121(b)
Leakage should be interpreted to include fuel discharged from the jet pipe under false start conditions, both on the ground and in flight.
See Orange Paper Amendment 96/1

ACJ 25.1125(a)(3)
Exhaust Heat Exchangers (Interpretative Material)
See JAR 25.1125(a)(3)
The cooling provisions should be arranged so that it is not possible to use the heat exchanger unless the cooling provisions are in operation.

ACJ 25.1141(f)
Powerplant Controls, General (Interpretative Material)
See JAR 25.1141(f)
A continuous indicator need not be provided.

ACJ 25.1181
Designated Fire Zones (Acceptable Means of Compliance)
See JAR 25.1181
[ 1 ISO/DIS 2685, (15 JULY 1992) "Aircraft - Environmental conditions and test procedures for airborne equipment - Resistance to fire in designated fire zones", gives test conditions and methods of demonstrating compliance with the "Fire-resistant " and "Fireproof " requirements. ]
2 Tests to demonstrate compliance with the standard grades of resistance to fire may not be necessary if similarity can be shown with other components which have been tested in accordance with this standard.
3 For example, materials which are considered satisfactory for use in firewalls without being subjected to fire tests include -
a. Stainless steel sheet 04 mm (0016 in) thick;
b. Mild steel sheet protected against corrosion 045 mm (0018 in) thick; and
c. Titanium sheet 045 mm (0018 in) thick.

ACJ 25.1195(b)
Fire Extinguisher Systems (Interpretative Material and Acceptable Means of Compliance)
See JAR 25.1195(b)
Acceptable methods to establish the adequacy of the fire extinguisher system are laid down in Advisory Circular 20-100.
ACJ - Subpart F

ACJ 25.1301(b)
Function and Installation (Acceptable Means of Compliance)
See JAR 25.1301(b)
1 Adequate means of identification should be provided for all cables, connectors and terminals. The means employed should be such as to ensure that the identification does not deteriorate under service conditions.
2 When pipelines are marked for the purpose of distinguishing their functions, the markings should be such that the risk of confusion by maintenance or servicing personnel will be minimised. Distinction by means of colour markings alone is not acceptable. The use of alphabetic or numerical symbols will be acceptable if recognition depends upon reference to a master key and any relation between symbol and function is carefully avoided. Specification ISO.12 gives acceptable graphical markings.

ACJ 25.1303(b)(5)
Altitude Displays (Interpretative Material and Acceptable Means of Compliance)
See JAR 25.1303(b)(5)
See Orange Paper Amendment 96/1
1 Attitude Displays (Interpretative Material)
1.1 For turbo-jet aeroplanes each display should be usable over the full range of 360 in pitch and in roll. For propeller-driven aeroplanes the pitch range may be reduced to 75 provided that no misleading indication is given when the limiting attitude is exceeded.
1.2 Paragraph 1.1 is not intended to prohibit the use of vertical references having controlled gyro precession, or its equivalent in the case of a stable platform, but precession should not occur at a pitch attitude closer to the horizontal than 70, and should be completed within an attitude change of 15.
1.3 The display should take the form of an artificial horizon line which moves relative to a fixed reference aeroplane symbol so as to indicate the position of the true horizon.
NOTES:
1 It is acceptable for the fixed reference aeroplane symbol to be positioned so that it is aligned with the horizon line during cruising flight.
2 If a variable index is provided in addition to the fixed aeroplane symbol it should be so designed that it will not introduce any risk of misinterpretation of the display.
[ 1.4 There should be no means accessible to the flight crew of adjusting the relationship between the horizon line and the reference aeroplane symbol. ]
[ 1.5 ] The artificial horizon line should move in roll so as to remain parallel to the true horizon, i.e. when the aeroplane rolls through an angle of 30 the artificial horizon line should also rotate through 30 relative to the fixed index.
[ 1.6 ] The artificial horizon line should remain in view over a range of pitch attitudes sufficient to cover all normal operation of the aeroplane plus a margin of not less than 2 in either direction. Additional 'ghost' horizon lines should be provided parallel to the main horizon line so that beyond this range at least one such line is in view at an attitude with the range of the display.
[ 1.7 ] The pitch attitude scale should be sensibly linear while the main horizontal line is in view, but may become non-linear beyond this range.
[ 1.8 All the attitude displays in the aeroplane should have a similar presentation so as to prevent any risk of confusion in transferring attention from one display to another. ]
1.9 Sufficient pitch and bank angle graduations and markings should be provided to allow an acceptably accurate reading of attitude and to minimise the possibility of confusion at extreme attitudes.
1.10 A bank angle index and scale should be provided. The index may be on the fixed or moving part of the display.
1.11 The 'earth' and 'sky' areas of the display should be of contrasting colours or shades. The distinction should not be lost at any pitch or roll angle.
1.12 Any additional information (e.g. flight director commands) displayed on an attitude display should not obscure or significantly degrade the attitude information.
1.13 The display should be clearly visible under all conditions of daylight and artificial lighting.
[ 1.14 Words which may be ambiguous (e.g. 'climb', 'dive', 'push', 'pull') should not be used. ]
2 Attitude Display Systems (Acceptable Means of Compliance)
2.1 The probability of indication of dangerously incorrect information without a warning being given should be Extremely Remote.
2.2 The warning may be provided by means of self- or comparison-monitoring and should be clear and unambiguous, e.g. a flashing light. Instrument flags are unlikely to be acceptable as a comparator warning unless they exclude a significant portion of the display in which case means should be provided to permit the removal of the flag from the display which is not in error.
2.3 The definition of dangerously incorrect information depends to some extent on the characteristics of the aeroplane, but in general an error greater than 5 in pitch or 10 in roll will be considered to be dangerous.

ACJ 25.1303(c)(1)
Flight and Navigation Instruments (Interpretative Material)
See JAR 25.1303(c)(1)
In the absence of warning through the inherent aerodynamic qualities of the aeroplane (e.g. buffeting) it should be shown that no single faults can result both in misleading airspeed information and in operation of the warning system outside its tolerances, such as would be likely to lead to exceedance of VMO/MMO.

ACJ 25.1305(d)(1)
Powerplant Instruments (Acceptable Means of Compliance)
See JAR 25.1305(d)(1)
The following are examples of parameters which are considered to be directly related to thrust; fan RPM(N1), integrated engine pressure ratio (IEPR) and engine pressure ratio (EPR), depending on engine type.
ACJ No. 2 to JAR 25.1309
Equipment, Systems and Installations (Interpretative Material)
See JAR 25.1309(a)
The effects of fluid or vapour contamination, due either to the normal environment or accidental leaks or spillage, should be taken into account.
ACJ No. 3 to JAR 25.1309
Equipment, Systems and Installations (Interpretative Material)
See JAR 25.1309(b)
The effects of mechanical damage or deterioration including short circuits or earths caused by such damage, in particular the failure of an earth connection should be taken into account.
ACJ No. 4 to JAR 25.1309
Equipment, Systems and Installations (Interpretative Material)
See JAR 25.1309(c)
Each source of electrical supply (e.g. generators and batteries) should be provided with means to give the flight crew immediate warning of the failure of its output. These warning means are additional to the system indication requirements of JAR 25.1351(b)(6). For multiphase systems the warning should also indicate the loss of any phase.
ACJ No. 6 to JAR 25.1309
Equipment, Systems and Installations (Acceptable Means of Compliance)
See JAR 25.1309(e)
Where alternative or multiplication of systems and equipment is provided to meet the requirements of JAR 25.1309(e), the segregation between circuits should be such as to minimise the risk of a single occurrence causing multiple failures of circuits or power supplies of the system concerned. For example, electrical cable bundles or groups of hydraulic pipes should be so segregated as to prevent damage to the main and alternative systems and power supplies.
ACJ No. 7 to JAR 25.1309
Equipment, Systems and Installations (Interpretative Material)
See JAR 25.1309(e)(3)
For aeroplanes for which the two-power-units-inoperative performance is scheduled, such services should remain operative as will enable the flight to be safely continued and terminated. In achieving this -
a. Some reduction in the performance of particular services is permissible (e.g. airframe ice-protection),
b. It may be assumed that electrical loads are reduced in accordance with a pre-determined procedure which is consistent with safety in the types of operation for which the aeroplane is certificated, and
c. Consideration should be given to any restrictions that may be necessary should the air supply for cabin pressure be interrupted or seriously reduced consequent upon the failure of the power-units.
ACJ No. 8 to JAR 25.1309
Equipment, Systems and Installations (Interpretative Material)
See JAR 25.1309(c)
1 The reliability of each warning system should be compatible with the general reliability of the system for which it provides a warning.
2 Each warning system should be designed so as to minimise unnecessary warnings.

ACJ 25.1315
Negative Accelerations (Acceptable Means of Compliance)
See JAR 25X1315
1 Demonstration of compliance with JAR 25X1315 should be made by analysis and/or ground tests, and should be supported by flight tests.
2 Analysis and/or Ground Tests. Appropriate analysis and/or ground tests should be made on components of essential fluid systems and such other components as are likely to be adversely affected by negative acceleration to demonstrate that they will not produce a hazardous malfunction.
3 Flight Tests
3.1 The aeroplane should be subjected to -
a. One continuous period of at least five seconds at less than zero g, and, separately,
b. A period containing at least two excursions to less than zero g in rapid succession, in which the total time at less than zero g is at least five seconds.
3.2 The tests should be made at the most critical condition from the fuel flow standpoint, e.g. with fuel flow corresponding to maximum continuous power and with the fuel representing a typical operational low fuel condition as for a missed approach.

ACJ 25.1321(a)
Instruments; Arrangement and Visibility (Interpretative Material)
See JAR 25.1321(a)
Where an optimum position for both pilots is not possible, any bias should be in favour of the first pilot.

ACJ 25.1323(c)(2)
Airspeed Indicating System (Interpretative Material)
See JAR 25.1323(c)(2)
From 13 VS to stall warning speed the rate of change of IAS with CAS should not be less than 075.

ACJ 25.1323(c)(3)
Airspeed Indicating System (Interpretative Material)
See JAR 25.1323(c)(3)
From VMO to VMO + (VDF - VMO) the rate of change of IAS with CAS should not be less than 05.

ACJ 25.1323(d)
Airspeed Indicating System (Interpretative Material)
See JAR 25.1323(d)
The design and installation of the pitot system should be such that positive drainage of moisture is provided, chafing of the tubing and excessive distortion at bends is avoided, and the lag and the possibility of moisture blockage in the tubing should be kept to an acceptable minimum.

ACJ 25.1323(e)
Airspeed Indicating System (Acceptable Means of Compliance)
See JAR 25.1323(e) and JAR 25.1325(b)
1 Tests should be conducted to the same standard as recommended for turbine engine air intakes (see ACJ 25.1093(b)(1)) unless it can be shown that the items are so designed and located as not to be susceptible to icing conditions. Ice crystal and mixed ice and water cloud will need to be considered where the system is likely to be susceptible to such conditions.
2 However, in conducting these tests due regard should be given to the presence of the aeroplane and its effect on the local concentration of the cloud.

ACJ 25.1325
Static Pressure System (Interpretative Material)
See JAR 25.1325
The lag and the possibility of moisture blockage in the tubing should be kept to an acceptable minimum.

ACJ 25.1328
Direction Indicator (Interpretative Material)
See JAR 25X1328
1 After correction the deviation on any heading should not exceed 1, except that -
a. On aeroplanes with a short cruising range, the above limit may be extended after consultation with the National Authority.
b. A change in deviation due to the current flow in any item of electrical equipment and its associated wiring is permissible, but should not exceed 1. The combined change for all such equipment, with all combinations of electrical load, should not exceed 2.
c. A change in deviation due to the movement of any component, (e.g. controls or undercarriage) in normal flight is permissible, but should not exceed 1.
2 The change in deviation due to the proximity of any item of equipment containing magnetic material should not exceed 1, and the combined change for all such equipment should not exceed 2.

ACJ 25.1329
Automatic Pilot (Interpretative Material and Acceptable Means of Compliance)
See JAR 25.1329
See Orange Paper Amendment 96/1
1 General
1.1 For the purpose of this ACJ the term 'automatic pilot' includes the sensors, computers, power supplies, servo-motors/actuators and associated wiring, necessary for its function. It includes any indications and controllers necessary for the pilot to manage and supervise the system.
1.2 Any part of the automatic pilot which remains connected to the primary flight controls when the automatic pilot is not in use is regarded as a part of the primary flight controls and the provisions for such systems are applicable.
1.3 In showing compliance with JAR 25.395(b), servo-motors, their mountings and their connection to the flight control system should have limit and ultimate factors of safety of not less than 10 and 15 respectively, with the maximum loads which can be imposed by the automatic pilot, or by the flight control system (up to its design load).
1.4 Adequate precautions should be taken in the design process and adequate procedures should be specified in the maintenance manual to prevent the incorrect installation, connection or adjustment of parts of the automatic pilot if such errors would hazard the aeroplane (e.g. torque clutches or limit switches with a range of adjustment such that maladjustment could be hazardous).
1.5 The response of the automatic pilot should be considered in showing compliance with the structural requirements of JAR-25 Subparts C and D.
1.6 The automatic pilot should be so designed and installed that the tolerances demonstrated during certification tests can be maintained in service.
1.7 The automatic pilot should not cause sustained nuisance oscillations, undue control activity or sudden large attitude changes, especially when configuration or power changes are taking place.
1.8 When automatic functions are provided which may be used with the automatic pilot (e.g. automatic throttle control or yaw damper, etc.) and use of the automatic pilot is permitted with any of these functions inoperative, it should comply with the provisions of this ACJ with these functions operative and inoperative.
1.9 Operating procedures for use with the automatic pilot should be established. (See JAR 25.1585(a) and (b).)
1.10 In addition to the quick release controls of JAR 25.1329(d), in order to show compliance with JAR 25.1309 an alternative means of disengagement, readily accessible in flight, should be provided.
1.11 It should be possible to disengage the automatic pilot at any time without unacceptable out-of-trim forces.
2 Performance of Function
2.1 The automatic pilot should be demonstrated to perform its intended function in all configurations in which it may be used throughout all appropriate manoeuvres and environmental conditions, including turbulence, unless an appropriate operating limitation or statement is included in the aeroplane Flight Manual. All manoeuvres should be accomplished smoothly, accurately and without sustained nuisance oscillation. This demonstration should be conducted with system tolerances at the lower limits of automatic pilot authority.
NOTE: The acceptability of the performance may be based on subjective judgement taking into account the experience acquired from similar equipment and the general behaviour of the aeroplane. The acceptable performance may vary according to aeroplane type and model.
2.2 If the automatic pilot is to be approved for ILS coupled approaches, a series of approaches should be made in the normal approach configuration(s) to the Minimum Use Height (MUH) (see paragraph 5.3.4). These approaches should be made in conditions chosen to show that the performance is satisfactory within permitted extremes such as weight, centre of gravity position, wind speed, capture angle and range, and more than one ILS beam should be used. To cover this range of conditions, it can be expected that in the order of 15 approaches will be needed. In the event that the performance is not satisfactory down to the MUH established in accordance with paragraph 5.3.4, then the Flight Manual should specify an MUH at which the performance is satisfactory. (An approach is considered to be satisfactory if it is stable without large deviations from the intended path or speed during the approach and, at the MUH, the position and velocities of the aeroplane are such that a safe landing can readily be made.)
See Orange Paper Amendment 96/1
2.3 If approach is sought for ILS coupled approaches initiated with one engine inoperative and the aeroplane trimmed at glide path intercept, the automatic pilot should be capable of conducting the approach without further manual trimming.
3 Controls, Indicators and Warnings
3.1 The controls, indicators and warnings should be so designed as to minimise crew errors. Mode and malfunction indications should be presented in a manner compatible with the procedures and assigned tasks of the flight crew. The indications should be grouped in a logical and consistent manner and be visible from each pilot's station under all expected lighting conditions.
3.2 The means provided to comply with JAR 25.1329(h) should also give an appropriate indication when there is -
a. Failure to achieve the selected mode; and
b. Inadvertent change or disengagement of a mode.
4 Characteristics of Some Specific Modes
4.1 Automatic Acquisition of Altitude Hold Mode. Where the automatic pilot has the ability to acquire and maintain a pre-selected altitude it should be shown in particular that -
a. If the pilot fails to advance the throttles following an altitude acquisition from a descent, the aeroplane exhibits no hazardous characteristics if recovery action is taken within a reasonable period after the onset of stall warning, or other appropriate warning;
NOTE: Compliance with this provision need not be demonstrated if adequate means are provided to prevent such an error.
b. Resetting the datum pressure or the selected altitude at any time during altitude acquisition does not result in a hazardous manoeuvre.
4.2 Go-around Mode. Where the automatic pilot has the ability to carry out an automatic go-around -
a. The speed should be compatible with that used for a manually controlled go-around; it should not be less than the higher of 12 VS or the appropriate minimum control speed (see JAR 25.149);
b. The control actions and flight path during the initial rotation should not be significantly different from those of a manually controlled go-around;
c. Flight path control following an engine failure during go-around should not require exceptional piloting skill or alertness; and
d. Any failure condition that causes the automatic pilot to fail to initiate the go-around without a warning appropriate to the approved use of the system, should be assessed as Extremely Remote.
4.3 Control Wheel Steering Mode (CWS). Where the pilot has the ability to make inputs to the automatic pilot by movement of the normal control wheel (control wheel steering) -
a. It should be possible for the pilot to overpower the automatic pilot and to achieve the maximum available control surface deflection without using forces so high that the controllability requirements of JAR 25.143(c) are not met;
b. The maximum bank and pitch attitudes which can be achieved without overpowering the automatic pilot should be limited to those necessary for the normal operation of the aeroplane;
NOTE: Typically 35 in roll +20 to -10 in pitch.
See Orange Paper Amendment 96/1
c. It should be possible to carry out all normal manoeuvres and to counter all normal changes of trim due to change of configuration or power, within the range of flight conditions in which control wheel steering may be used, without encountering excessive discontinuities in control force which might adversely affect the flight path;
d. The stall and stall recovery characteristics of the aeroplane should remain acceptable. It should be assumed that recovery is made with CWS in use unless automatic disengagement of the automatic pilot is provided;
e. In showing compliance with JAR 25.143(f) account should be taken of such adjustments to trim as may be carried out by the automatic pilot in the course of manoeuvres which can reasonably be expected. Some alleviation may be acceptable in the case of unusually prolonged manoeuvres provided the reduced control forces would not be hazardous;
f. If the use of this mode for take-off and landing is to be permitted it should be shown that -
i. Sufficient control, both in amplitude and rate is available without encountering force discontinuities;
ii. Reasonable mishandling is not hazardous (e.g. engaging the automatic pilot while the elevators or ailerons are held in an out-of-trim position); and
iii. Runaway rates and control forces are such that the pilot can readily overpower the automatic pilot with no significant deviation in flight path;
See Orange Paper Amendment 96/1
See Orange Paper Amendment 96/1
g. It should not be possible to engage CWS by applying a force to the control column or wheel when the automatic pilot is coupled to an ILS localiser and glide path.
5 Failure Conditions
5.1 Analysis
5.1.1 An analysis should be carried out to define the Failure Conditions and their Effects and to show that the probability of each Failure Condition is such that the provisions of paragraph 5.2 are achieved. The depth of the analysis may be significantly reduced and numerical probability analysis may not be required in the case of a single-channel automatic pilot if worst-case failures can be easily identified and used as the basis of a ground and flight test demonstration programme (e.g. where the effect of a failure is limited by an independent device whose serviceability is frequently checked).
5.1.2 When the failure of a device can remain undetected in normal operation, the frequency with which the device is checked will directly influence the probability that such a failure is present on any particular occasion. This should be taken into account when assessing the probabilities of any Failure Conditions which include dormant failures in the monitoring devices or in other unchecked parts of the system (see paragraph 5.1.6).
5.1.3 When the failure of a component or equipment can be expected to result in other failures, then these further failures should be taken into account in the analysis. In assessing which further failures may occur, consideration should be given to any change in the equipment operating conditions for other components or equipment resulting from the first failure.
5.1.4 In considering damage from external sources, account should be taken of the location of the equipment in the aeroplane and other features of the installation.
5.1.5 Attention should be given in the analysis to common mode failures (i.e. multiple failures arising from a single cause). The following are examples:
a. A local fire causing multiple fractures;
b. Electromagnetic interference or electrical transients causing multiple malfunctions;
c. Mechanical vibration causing multiple failures or malfunctions;
d. Leakage of water or other liquids (e.g. from galley, lavatories or cargo) causing multiple electrical failures;
e. The failure of a cooling system or the leakage of hot air causing multiple failures in other systems;
f. Lightning strike; and
g. Engine failure.
5.1.6 When exposure times relevant to failure probability calculations are dependent on flight crew and maintenance checks (i.e. pre-flight, first flight of the day, pre-land etc.) and/or inspection intervals for dormant (latent) failures, these tasks, time intervals and the recommended component monitoring programme should be clearly specified in the certification documentation, and made available for the purposes of scheduling flight crew and maintenance procedures.
5.1.7 For digital systems and software development, verification, testing and system level validation processes should be accomplished in accordance with an appropriate methodology. The RTCA Document DO 178A or EUROCAE ED 12A contain guidance as to methodologies which may be used, at the appropriate levels, to establish compliance. Other means of establishing compliance may be acceptable.
5.2 Acceptability of Failure Conditions
5.2.1 Any Failure Condition occurring within the normal flight envelope should be assessed as Extremely Improbable if its effect is one of the following:
a. A load on any part of the primary structure sufficient to cause a catastrophic structural failure;
b. Catastrophic loss of flight path control;
c. Exceedance of VDF/MDF; or
d. Catastrophic flutter or vibration.
5.2.2 Any Failure Condition occurring within the normal flight envelope should be assessed as Extremely Remote if its effect is one of the following:
a. A load on any part of the structure greater than its limit load;
b. Exceedance of an airspeed halfway between VMO and VDF or a Mach number halfway between MMO and MDF;
c. A stall;
d. A normal acceleration less than a value of 0 g;
See Orange Paper Amendment 96/1
e. Bank angles of more than 60 en route or more than 30 below a height of 1000 ft (3048 m);
f. Hazardous degradation of the flying qualities of the aeroplane;
g. Hazardous height loss in relation to minimum permitted height for automatic pilot use (see paragraph 5.3); or
h. Engagement or disengagement of a mode leading to hazardous consequences.
5.2.3 Any Failure Condition for which the probability of occurrence is assessed as Remote should have an appropriately less severe effect than those listed in paragraph 5.2.2.
5.2.4 Compliance with the requirements of paragraphs 5.2.1, 5.2.2 and 5.2.3 should be shown by ground simulation, flight tests or suitable analysis. Where appropriate, account should be taken of pilot recognition of the Failure Condition, and any subsequent recovery action taken. The limiting values given in paragraph 5.2.2 should not be exceeded either during any manoeuvre caused by the failure or during the recovery by the pilot. The minimum heights at which the automatic pilot may be used should be determined.
5.2.5 The most critical of the Failure Conditions which are not assessed as Extremely Remote or Extremely Improbable should be demonstrated in flight test (see paragraph 5.3). Failure Conditions which are assessed as Extremely Remote may be demonstrated by a ground simulation or analysis which has been suitably validated, using the same procedures as are specified in paragraph 5.3 for flight test.
5.3 Flight Demonstrations. When demonstrating compliance with paragraph 5.2 by means of flight test, the following procedures should be used:
5.3.1 General
a. Failure Conditions of the automatic pilot including, where appropriate, multi-axis failures and automatic-trim failures, should be simulated in such a manner as to represent the overall effect of each Failure Condition about all axes.
See Orange Paper Amendment 96/1
b. Following recognition of the Failure Condition by the pilot, a delay, as specified in paragraphs 5.3.2, 5.3.3, 5.3.4 and 5.3.5 should be applied before the commencement of recovery action. Following such delay the pilot should be able to return the aeroplane to its normal flight attitude under full manual control without engaging in any dangerous manoeuvres during recovery and without control forces exceeding the values given in JAR 25.143(c). During the recovery the pilot may overpower the automatic pilot or disengage it. For the purpose of determining the minimum height at which the automatic pilot may be used during an approach the pilot should not apply a normal acceleration exceeding 15 g (total).
See Orange Paper Amendment 96/1
c. System authority should be set at the most adverse tolerance limits, and the flight condition should be the most critical which is appropriate (centre of gravity, weight, flap setting, altitude, speed, power or thrust).
d. In malfunction tests described in paragraphs 5.3.2, 5.3.3, 5.3.4 and 5.3.5 the recognition point should be that at which a pilot in service operation in non-visual conditions may be expected to recognise the need to take action and not that at which the test pilot engaged in the flight trials does so. Recognition of the malfunction may be through the behaviour of the aeroplane or an appropriate failure warning system and the recognition point should be identified. Control column or wheel movements alone should not be used for recognition. The recognition time should not normally be less than 1 second. If a recognition time of less than 1 second is claimed, specific justification will be required (e.g. additional tests to ensure that the time is representative in the light of the cues available to the pilot).
See Orange Paper Amendment 96/1
e. If automatic throttles are installed the tests should be carried out with the automatic throttles operating, and not operating.
f. For control wheel steering, in those phases of flight where the pilot is exercising manual control (e.g. take-off, landing) the delay times specified in paragraphs 5.3.2, 5.3.3, and 5.3.5 need not be applied. The pilot may commence recovery action at the recognition point. (See also paragraph 4.3 f.)
g. The aeroplane should be so instrumented that the parameters appropriate to the test are recorded (e.g. normal acceleration, airspeed, height, pitch and roll angles, automatic pilot engagement state). The fitment of the instrumentation should not affect the behaviour of the automatic pilot or any other system.
5.3.2 Climb, Cruise, Descent and Holding
a. Recovery action should not be initiated until three seconds after the recognition point.
b. The MUH for the automatic pilot in climb, cruise, descent or holding should not be less than 1000 ft (3048 m), unless the height loss is determined under the conditions for which use of the automatic pilot is requested. In that case the MUH should not be less than twice the height loss. The height loss is measured as the difference between the height at the time the malfunction is induced to the lowest height in the recovery manoeuvre.
5.3.3 Manoeuvring Flight
a. Recovery action should not be initiated until one second after the recognition point.
b. Malfunctions should be induced in turns at the maximum bank angles for normal operation.
5.3.4 Approach Coupled to an ILS Glide Path
a. The aircraft should be flown down the ILS in the configuration and at the approach speed specified by the applicant for approach. Simulated automatic pilot malfunctions should be induced at critical points along the ILS, taking into consideration all possible design variations in automatic pilot system sensitivity and authority. In general, malfunction demonstrations may be restricted to hardovers (and possibly automatic-trim failures) unless an MUH below 100 ft (3048 m) is requested, when runaways at lower rates should also be investigated.
b. A 3 glide path should be used.
c. The aeroplane should be so instrumented that the following information is recorded:
i. The path of the aeroplane with respect to the normal glide path;
ii. The point along the glide path when the simulated malfunction is induced;
iii. The point where the pilot indicates recognition of the malfunction; and
iv. The point along the path of the aeroplane where the recovery action is initiated.
d. Recoveries from malfunction should simulate non-visual conditions with a one-second time delay between recognition point and initiation of recovery.
e. The MUH should be determined as the height of the aeroplane wheels at the point where recovery from the failure is initiated when the path of the aeroplane wheels during the recovery manoeuvre is tangent to the runway or to a 1:29 slope line drawn from a point 15 ft (457 m) above the runway threshold (See Figure 1). If there is no automatic landing capability, the MUH should not be less than 50 ft (1524 m).
See Orange Paper Amendment 96/1
f. An engine failure during an ILS coupled approach should not cause a heading change at a rate greater than three degrees per second or produce hazardous attitudes (see also paragraph 5.2.2 e.). In showing compliance with this, manual retrimming of the aeroplane is not permitted.
5.3.5 Approach not coupled to an ILS Glide Path
a. The procedure described in paragraphs 5.3.4 a. to f. should be applied.
b. A descent path of three degrees should be used unless the automatic pilot is to be approved for significantly steeper descents.
c. The MUH for the automatic pilot should not be less than twice the height loss, where the height loss is measured as described in paragraph 5.3.2 b.
5.3.6 Failure to disengage. Unless failure of the automatic pilot to disengage during the approach when the pilot operates the quick release control on the control wheel is assessed as Extremely Remote it should be demonstrated that the pilot can control the aeroplane manually without operating any of the other disengagement controls.
See Orange Paper Amendment 96/1
5.4 Oscillatory tests
5.4.1 An investigation should be made to determine the effects of an oscillatory signal of sufficient amplitude to saturate the servo amplifier of each device that can move a control surface unless such a malfunction is assessed as Extremely Improbable. The investigation should cover the range of frequencies which can be induced by a malfunction of the automatic pilot and systems functionally connected to it, including an open circuit in a feed-back loop. The investigated frequency range should include the highest frequency which results in apparent movement of the system driving the control surface to the lowest elastic or rigid body response frequency of the aeroplane. Frequencies less than 02 cps may, however, be excluded from consideration. The investigation should also cover the normal speed and configuration ranges of the aeroplane. The results of this investigation should show that the peak loads imposed on the parts of the aeroplane by the application of the oscillatory signal are within the limit loads for these parts.
5.4.2 The investigation may be accomplished largely through analysis with sufficient flight data to verify the analytical studies or largely through flight tests with analytical studies extending the flight data to the conditions which impose the highest percentage of limit load to the parts.
5.4.3 When flight tests are conducted in which the signal frequency is continuously swept through a range, the rate of frequency change should be slow enough to permit determining the amplitude of response of any part under steady frequency oscillation at any critical frequency within the test range
See Orange Paper Amendment 96/1
6 Aeroplane Flight Manual. The aeroplane Flight Manual should contain the following:
a. Any approved limits on the use of the automatic pilot necessary for compliance with the provisions of paragraphs 1 to 5 (e.g. aeroplane configuration, speed, altitude, wind, temperature, and MUH for the automatic pilot);
See Orange Paper Amendment 96/1
NOTE: Use of the automatic pilot with low weather minima may vary the items listed above. The aeroplane Flight Manual should contain the appropriate information.
b. Procedures necessary for compliance with the provisions of paragraphs 1 to 5 (e.g. safety checks, abnormal and emergency procedures, and operation in wind shear and turbulence etc.)

FIGURE 1 - DEVIATION PROFILE METHOD FOR DETERMINING MINIMUM
HEIGHT AT WHICH THE AUTOMATIC PILOT MAY BE USED (MUH)
ON AN ILS COUPLED APPROACH


ACJ 25.1331(a)(3)
Instruments Using a Power Supply (Interpretative Material)
See JAR 25.1331(a)(3)
Where practicable, the warning should be incorporated in the instrument.

ACJ 25.1333(b)
Instrument Systems (Interpretative Material and Acceptable Means of Compliance)
See JAR 25.1333(b)
1 In showing compliance with JAR 25.1333(b) account may be taken of the probability with which loss of information will lead to a catastrophe.
2 Attitude Display Systems. One acceptable means of compliance with JAR 25.1333(b) is to provide three displays, the reliability and independence of which should be confirmed by a suitable assessment. Each display should have independent sensors and power supplies. The power supply to the standby display and its lighting should be such that the display is usable for not less than 30 minutes if a total failure of the generated electrical power causes the loss of both main instruments.
NOTE: The time for which the display remains usable will be stated in the Flight Manual.

ACJ 25.1351(b)(5)
[ Generating System (Acceptable Means of Compliance and Interpretative Material) ]
See JAR 25.1351(b)(5)
[ 1 ] The disconnect means required by JAR 25.1351(b)(5) should be accessible to the appropriate flight-crew members in their normal seated positions.
[ 2 The power source controls should be considered as cockpit controls and therefore also comply with JAR 25.777.
3 It may not be necessary to provide disconnection controls for all power sources, for example RAT generators or engine control dedicated generators. Where it is necessary to isolate the alternate power source when normal generator power is restored, such isolation should be possible. ]

ACJ 25.1351(d)
[ Operation without Normal Electrical Power (Interpretative Material)
See JAR 25.1351(d)
1 Provision should be made to ensure adequate electrical supplies to those services which are necessary to complete the flight and make a safe landing in the event of a failure of all normal generated electrical power. All components and wiring of the alternate supplies should be physically and electrically segregated from the normal system and be such that no single failure, including the effects of fire, the cutting of a cable bundle, the loss of a junction box or control panel, will affect both normal and alternate supplies.
2 When ensuring the adequacy of electrical supplies relative to alternate power source duration and integrity, special consideration should be given to aeroplanes such as those with fly-by-wire, for which the total loss of electrical supplies could result in an immediate loss of control.
3 In considering the services which should remain available following the loss of the normal generated electrical power systems, consideration should be given to the role and flight conditions of the aeroplane and the possible duration of flight time to reach an airfield and make a safe landing.
4 The services required by JAR 25.1351(d)(1) may differ between aeroplane types and roles and should be agreed with the Authority. These should normally include - ]
[ a. Attitude information;
b. Radio communication and intercommunication;
c. Navigation;
d. Cockpit and instrument lighting;
e. Heading, airspeed and altitude, including appropriate pitot head heating;
f. Adequate flight controls;
g. Adequate engine control; and
Restart capability with critical type fuel (from the standpoint of flame-out and restart capability) and with the aeroplane initially at the maximum certificated altitude;
h. Adequate engine instrumentation;
i. Such warning, cautions and indications as are required for continued safe flight and landing;
j. Any other services required for continued safe flight and landing.
5 Consideration should also be given to the equipment and the duration of services required to make a controlled descent and forced landing in the event of failure and inability to restart all engines.
6 Alternate Power Source Duration and Integrity
6.1 Time Limited. Where an alternate power source provided to comply with JAR 25.1351(d) is time limited (e.g. battery), the required duration will depend on the type and role of the aeroplane. Unless it can be shown that a lesser time is adequate, such a power source should have an endurance of at least 60 minutes, at least 30 minutes of which is available under IMC. An endurance of less than 30 minutes under IMC would not normally be acceptable. The endurances, with any associated procedures, should be specified in the Flight Manual. The endurance time should be determined by calculation or test, due to allowance being made for -
a. Delays in flight crew recognition of failures and completion of the appropriate drill where flight crew action is necessary. This should be assumed to be 5 minutes provided that the failure warning system has clear and unambiguous attention-getting characteristics and where such a delay is acceptable and compatible with the crew's primary attention being given to other vital actions.
b. The minimum voltage acceptable for the required loads, the battery state of charge, the minimum capacity permitted during service life and the battery efficiency at the discharge rates and temperatures likely to be experienced. Unless otherwise agreed, for the purpose of this calculation, a battery capacity at normal ambient conditions of 80% of the nameplate rated capacity, at the one hour rate, and a 90% state of charge, may be assumed (i.e. 72% of nominal demonstrated rated capacity at +20C). The allowance for battery endurance presumes that adequate requirements for periodic battery maintenance have been agreed.
c. For those aeroplanes where the battery is also used for engine or APU starting on the ground, it should be shown that following engine starts, the charge rate of the battery is such that the battery is maintained in a state of charge that will ensure adequate alternate power source duration should a failure of generated power occur shortly after take-off.
NOTE: This may normally be achieved by ensuring that, following battery-powered starting, the battery charge current has fallen to a declared level prior to take-off.
d. For those aeroplanes where the battery is used for in-flight starting of the engines or APU, it may be necessary to include limitations on the number of attempted starts, or to provide a separate dedicated battery for such purposes. ]
[ 6.2 Non Time Limited. Where an alternate electrical supply is provided by a non time limited source, e.g. APU, ram air turbine, pneumatic or hydraulic motor, due account should be taken of any limitation imposed by aeroplane speed, attitude, altitude etc., which may affect the capabilities of that power source. In considering the power source, account should be taken of the following:
a. Auxiliary Power Unit (APU). An APU capable of continuous operation throughout an adequate flight envelope may be considered an acceptable means of supplying electrical power to the required services provided that its air start capability is adequate and may be guaranteed. Where, however the APU is dependent for its starting current on a battery source which is supplying critical loads, such starting loads may prejudice the time duration of the flight if APU start is not achieved.
It may be necessary therefore to include limitations on the number of attempted starts or to provide a separate battery for APU starting, if this method of supplying electrical power is adopted. Consideration should also be given to the equipment, services and duration required prior to the APU generator coming on-line. Common failures which could affect the operation of all engines and the APU should be taken into consideration, e.g. fuel supply.
b. Ram Air Turbine (RAT). A ram air turbine may be utilised to provide an alternate electrical power source, but due consideration must be given to ensuring that the means of bringing the unit into use are not dependent on a source which may have been lost as a result of the original failure. This will normally necessitate independent, duplicate means of deployment. Particular attention should be given to ensuring that the RAT and its means of deployment satisfy the overall reliability requirements.
The continuity of electrical power to those services which must remain operative without crew action prior to the RAT being brought into operation, may necessitate the use of a battery, unless the operation of the emergency power source is automatic and is supplying power within a timespan so as not to jeopardise the continued safety of the aeroplane in the event of failure of normal generated electrical power.
c. Pneumatic or Hydraulic Motor Drive Power Source. A pneumatic or hydraulic motor driven electrical power source may be utilised subject to the same constraints on activation as the ram air turbine (see 6.2(b)). Care should be taken in ensuring that the operation of the pneumatic or hydraulic system is not prejudiced by faults leading to, or resulting from, the original failure, including the loss of, or inability to restart all engines.
d. Regaining of Main Generators. In the event of a major loss of electrical power, provision may be made for regaining the output of one or more generators using separate control and switching arrangements on the generator side of the normal generator line contactor. Such a system would not normally be acceptable on aeroplanes with less than three engine-driven generators, as the probability of the loss of all engine-driven generators is unlikely to meet the requirements of JAR 25.1351(d). To comply with JAR 25.1351(d)(2) the system should be designed such that the loss of both the main and alternate means of control and distribution is Extremely Improbable. Consideration should be given to the services and duration required prior to the activation of the system and to enable a descent and forced landing to be made, in the event of the inability to restart all engines. ]

ACJ 25.1353(a)
Electrical Equipment and Installations (Interpretative Material)
See JAR 25.1353(a)
The possible sources of interference to be considered should include -
a. Conducted and radiated interference caused by electrical noise generation from apparatus connected to the busbars,
b. Coupling between electrical cables or between cables and aerial feeders,
c. Malfunctioning of electrically-powered apparatus,
d. Parasitic currents and voltages in the electrical distribution and earth systems, including the effects of lightning currents or static discharge,
e. Difference frequencies between generating or other systems, and
f. The requirements of JAR 25.1309 should also be satisfied.
[ ACJ 25.1353(c)(6)(ii) and (iii)
Electrical Equipment and Installations (Interpretative Material)
See JAR 25.1353(c)(6)(ii) and (iii)
Where temperature sensing and over-temperature warning devices are installed to comply with JAR 25.1353(c)(6)(ii) or (iii), their correct operations should be verified at agreed maintenance intervals in addition to compliance with JAR 25.1309(a) and (b). ]

ACJ 25.1355(c)
Distribution System (Interpretative Material)
See JAR 25.1355(c)
The arrangement, protection and control of the feeders from the busbars to the distribution points, and the divisions of loads among the feeders, should be such that no single fault occurring in any feeder or associated control circuit will hazard the aeroplane.

ACJ 25.1357(a)
Circuit Protective Devices (Interpretative Material)
See JAR 25.1357(a)
No hazard should result from the effects of variations in ambient temperatures on either the protective device or the equipment it protects. See also JAR 25.1309.

ACJ 25.1360(a)
Protection Against Injury (Acceptable Means of Compliance)
See JAR 25X1360(a)
1 Where there may be a hazard during maintenance or servicing, aeroplane panels, etc., carrying voltages of above 50V RMS, should be marked with the voltage.
2 Where socket outlets are provided, e.g. for electric razors, these should be labelled as to use and with the output voltage or voltages. Where the output voltage exceeds 100 volts d.c. and/or 50 volts a.c. RMS either the output should be electrically isolated from the aeroplane structure, or means shall be provided to prevent inadvertent contact with live parts.

ACJ 25.1360(b)
Protection Against Injury (Acceptable Means of Compliance)
See JAR 25X1360(b)
1 For equipment which has to be handled during normal operation by the flight or cabin crew, a temperature rise of the order of 25C, for metal parts, should not be exceeded. For other equipment, mounted in parts of the aeroplane normally accessible to passengers or crew, or which may come into contact with objects such as clothing or paper, the surface temperature should not exceed 100C, in an ambient temperature of 20C.
2 The heating surfaces of properly installed cooking apparatus are excluded from these requirements.
3 The provision of guards around hot surfaces is an acceptable method of complying with these requirements.
[ ACJ 25X1362
Electrical Supplies for Emergency Conditions (Interpretative Material)
See JAR 25X1362
1 Consideration should be given to the possibility that all electrical power sources are likely to be disconnected or isolated by the flight crew just prior to, or during, an emergency (or crash) landing, to prevent them becoming a source of ignition.
2 In order that it shall not be necessary to reconnect power sources to enable a power supply to be provided to the emergency services, it would be acceptable to power such services from a 'hot' battery bus. These circuits would need to be so protected that the risk of their causing a fire under these conditions is minimised.
3 The emergency services which may require such a supply should include fuel and hydraulic shut-off valves, engine and APU fire extinguisher systems. (See also JAR 25.1189 and 25.1195). ]
[ ACJ 25.1363
Electrical Systems Tests (Acceptable Means of Compliance)
See JAR 25.1363
1 In carrying out the tests due account should be taken of load switching and flight crew operation of the system.
2 Laboratory or Ground Tests
2.1 All tests should be carried out with all equipment as representative as possible of the actual aeroplane. In particular, the simulation should include the correct representation of aeroplane cables in size, length and impedance, the correct ground (airframe) impedance and relative ground plane location and their location to other cables or systems, that could influence performance. System loads and the generator drive system should also be correctly simulated.
2.2 The tests may be carried out on representative laboratory rigs or in an actual aeroplane, as appropriate.
2.3 Test procedures should be prepared to cover each test condition in the programme.
3 Aeroplane Flight Tests
3.1 If not adequately simulated by laboratory or ground testing, flight tests should be carried out as necessary.
3.2 Temperature tests should be carried out on equipment to establish the adequacy of the cooling media under all ground and flight conditions.
3.3 Measurements should be made to ensure that all equipment, particularly the aeroplane battery, is operating within its specified environmental conditions.
3.4 Test procedures should be prepared to cover the conditions of the tests. ]

ACJ 25.1419
Ice Protection (Interpretative Material and Acceptable Means of Compliance)
See JAR 25.1419
1 General. Two ways of showing compliance with JAR 25.1419 are given.
1.1 Method 1. Method 1 is an arbitrary empirical method based on United Kingdom and French practice. This method is acceptable to all participating countries.
1.2 Method 2. Method 2 is a general approach based on US practice in applying FAR Part 25, Appendix C. If this method is used, each application will have to be evaluated on its merits.
1.3 Additional material, based on UK practice, appropriate to operating in ice crystal conditions is given in 4. This material should be used only where design features of the aeroplane are susceptible to this form of icing.
2 Method 1 (Acceptable Means of Compliance)
2.1 Any part of the aeroplane (including its equipment) which is susceptible to ice accretion in ice forming conditions, should be subjected to such evaluation as would demonstrate the suitability of the aeroplane to fly in the ice forming conditions defined in JAR Appendix C.
2.2 For the purposes of analysis and tests on protected surfaces the conditions of Figures 1, 2, 4 and 5 only of Appendix C should apply. In determining the rates of catch, the full spectrum of the droplet sizes should be considered but in determining impingement areas, a maximum droplet size of 50 m need only be considered.
2.3 The natural icing tests carried out on the aeroplane will be judged for their acceptability by evaluation of icing conditions through which the aeroplane has flown in relation to the envelope of conditions of Appendix C.
2.4 Where there are parts of the aeroplane which are not amenable to analysis, or when testing is considered necessary, the following paragraphs describe an acceptable method of demonstration that the requirements are complied with.
2.5 Protected and Unprotected Parts of the Airframe
2.5.1 General. When considering simulated icing tests, the flight conditions selected for testing at each temperature should be the most unfavourable taking account of aeroplane speed, altitude, angle of incidence and power supply. Where altitude is a critical parameter, the tests should be conducted in flight or on the ground so as to simulate the effects of altitude. When the tests are conducted in non-altitude conditions the system supply and the external aerodynamic and atmospheric conditions should be so modified as to represent the required altitude conditions as closely as possible.
2.5.2 Tests in Continuous Maximum Conditions
a. Those parts of the airframe where the accretion of ice under the conditions of Appendix C is likely to have an adverse effect on the airworthiness of the aeroplane, should be tested for a period of 30 minutes duration at each of the conditions specified in the following Table 1.
TABLE 1
Atmospheric Temperature Liquid Water Content Mean effective drop
(C) (g/m3) diameter (m)
0 08
-10 06
-20 03 20
-30 02

b. At the end of the tests the total ice accretion should be such as not to adversely affect the safety of the aeroplane.
c. The duration of the above tests can be reduced if it can be demonstrated that the surface is completely ice free or that the total ice accretion is obviously contained by repetitive shedding either naturally or enforced by cyclic operation of the protective system.
2.5.3 Check Concerning Intermittent Maximum Conditions. It would be necessary to check that Intermittent Maximum icing conditions of Figures 4 and 5 of Appendix C do not hazard the aeroplane. The encounters considered should include three clouds of 5 km horizontal extent with Intermittent Maximum concentrations as in Table 2 separated by spaces of clear air of 5 km.
TABLE 2
Atmospheric Temperature Liquid Water Content Mean effective drop
(C) (g/m3) diameter (m)
0 25
-10 22
-20 17 20
-30 10

2.5.4 Ice Accretion on Unprotected Parts
a. Where ice can accrete on unprotected parts is should be demonstrated that the effect of such ice will not critically affect the characteristics of the aeroplane as regards safety (e.g. flight, structure and flutter). The subsequent operation of retractable devices should be considered.
b. Irrespective of what is required by paragraphs 2.5.2 and 2.5.3 from service experience the amount of ice on the most critical unprotected main aero-foil surface need not exceed a pinnacle height of 75 mm (3 in) in a plane in the direction of flight. For other unprotected main surfaces an analysis may be performed to determine the maximum ice accretion associated with this maximum pinnacle height. In the absence of such an acceptable analysis a uniform pinnacle height of 75 mm (3 in) should be assumed. The shape and apparent density, taking into account the texture of the ice, are important. Unless suitable evidence is already available, icing tests should be conducted to determine the critical values of these properties.
c. The critical ice accretion on unprotected parts will normally occur during the hold near 15 000 feet at about -10C so as to give a total temperature of around 0C.
2.5.5 Ice Shedding. Parts of the aeroplane which can accrete ice which upon shedding could interfere with the continuous safe operation of the engines or essential services should be so protected as to prevent the shedding of ice having more than critical dimensions for the engine or device or it should be demonstrated that the trajectories of such ice are not critical. The protection or otherwise should be demonstrated assuming the ice conditions against which the engine air intake is required to be demonstrated.
2.5.6 Essential Equipment. Tests should be conducted to the same standard as recommended for turbine engine air intakes (see ACJ 25.1093(b)(1)) unless it can be shown that the items are so designed and located as not to be susceptible to icing conditions. Ice crystal and mixed ice and water cloud will need to be considered. However, in conducting these tests due regard should be given to the presence of the aeroplane and its effect on the local concentration of the cloud.
3 Method 2 (Interpretative Material)
3.1 Any part of the aeroplane (including its equipment) which is susceptible to ice accretion in ice-forming conditions, should be subjected to such evaluations as would demonstrate the suitability of the aeroplane to fly in ice-forming conditions defined in JAR 25, Appendix C, using FAA Advisory Circular AC 20-73, dated 21st April, 1971, and FAA Technical Report ADS4, dated March, 1964.
3.2 Factors which should be considered in the evaluation are -
a. The meteorological conditions of Appendix C,
b. The operational conditions which would affect the accumulation of ice on protected and unprotected surfaces of the aeroplane,
c. The operational conditions of the engine and propeller (if applicable) which would affect the accumulation of ice and/or the availability of energy to operate systems, and
d. The local condition resulting from installation on the aeroplane.
3.3 For the purpose of analysis and tests on protected and unprotected surfaces, all Figures 1 to 6 of Appendix C are used. In determining the more critical conditions of rate of catch and limits of impingements, the full spectrum of droplet sizes should be considered, taking into account the droplet size distribution (Langmuir D distribution is acceptable for this use).
3.4 The natural icing tests carried out on the aeroplane will be judged for their acceptability by the evaluation of the icing conditions through which the aeroplane has flown in relation to the envelope of conditions of Appendix C.
3.5 In following the alternative procedures as listed in JAR 25.1419(c)(1) and (3), the conditions selected for testing should be the most critical as determined from the analysis.
3.6 Where ice can accrete on protected or unprotected parts it should be demonstrated that the effect of such ice will not critically affect the characteristics of the aeroplane as regards safety (e.g. flight, structure and flutter). The subsequent operation of retractable safety devices should be considered.
3.7 From service experience the amount of ice on the most critical unprotected main aerofoil surface need not usually exceed a pinnacle height of 75 mm (3 in) in a plane in the direction of flight. For other unprotected main surfaces an analysis may be performed to determine the maximum ice accretion associated with this maximum pinnacle height. In the absence of such an acceptable analysis a uniform pinnacle height of 75 m (3 in) should be assumed. The shape and apparent density, taking into account the texture of the ice, are important. Unless suitable evidence is already available, icing tests should be conducted to determine the critical values of these properties.
3.8 The critical ice accretion on unprotected parts will normally occur during the hold near 15 000 feet so as to give a total temperature of around 0C.
3.9 Parts of the aeroplane which can accrete ice, which, upon shedding, could interfere with the continuous safe operation of the engines or essential services should, if necessary, be so protected as to prevent the shedding of ice having more than critical dimensions for the engine or device, or it should be demonstrated that the trajectories of such ice are not critical.
4 Ice Crystal Conditions. An assessment should be made into the vulnerability of the aeroplane and its systems to ice crystal conditions.
4.1 The parts most likely to be vulnerable are -
a. Turbine engine intakes with bends, particularly reverse flow (see JAR 25.1093), and
b. Pitot heads, etc. (see JAR 25.1323 and 1325).
4.2 Other parts requiring evaluation could be -
a. Ducts supplying essential air e.g. cooling, and
b. APU intakes (see ACJ 25.1093(b)(2)).
4.3 Where any doubt exists as to the safe operation in ice crystal conditions appropriate tests should be conducted to establish the proper functioning of the system likely to be affected.
4.4 For guidance Table 3 gives provisional details of the conditions likely to be encountered in service.
TABLE 3
Air Temperature (C) Altitude Range Maximum Horizontal Extend Mean
Crystal Content Particle
Diameter
(ft) (m) (g/m3) (km) (n miles) (mm)
0 to -20 10 000 3000 50 5 (3)
to to 20 100 (50)
30 000 9000 10 500 (300)
________________________________________________________________________ 1.0
-20 to -40 15 000 4500 50 5 (3)
to to 20 20 (10)
40 000 12 000 10 100 (50)
05 500 (300)
NOTES:
1 In the temperature range 0 to -10C the ice crystals are likely to be mixed with water droplets (with a maximum diameter of 2 mm) up to a content of 1 g/m3 or half the total content whichever is the lesser, the total content remaining numerically the same.
2 The source of information is RAE Tech, Note Mech. Eng. 283 dated May 1959.

ACJ 25.1435(a)(4)
Hydraulic Systems (Interpretative Material)
See JAR 25.1435(a)(4) and (a)(7)
It is recommended that, in achieving compliance with this requirement, reliance should not be placed upon a simple pressure relief device. Experience gained from hydraulic systems in which the pump has failed to off-load and has therefore delivered maximum flow at maximum pressure, shows that the resultant temperature rise across the pressure-reducing valve can produce fluid degradation and a potentially serious fire hazard, depending on the type of fluid being used. This may also affect the integrity of items such as joints, seals and flexible hoses.

ACJ 25.1435(a)(8)
Hydraulic Systems (Acceptable Means of Compliance)
See JAR 25.1435(a)(8)
Prevention of hazard may be achieved either by design of the pump or by its location or by both.

ACJ 25.1435(b)(2)
Hydraulic Systems (Interpretative Material)
See JAR 25.1435(b)(2)
The loads due to vibration and the loads due to temperature effects are those loads which act upon the elements of the system due to environmental conditions.

ACJ 25.1436(b)(3)
Pneumatic Systems (Interpretative Material)
See JAR 25X1436(b)(3)
1 In systems in which the air pressure of the supply sources is significantly greater than the system operating pressure (e.g. an engine bleed-air tapping) due account should be taken of the consequences of failure of the pressure-regulating device when assessing the strength of the system, downstream of the device relative to the values of PW, PL and PR.
2 Such devices should be protected as necessary against deleterious effects resulting from the presence of oil, water or other impurities which may exist in the system.

ACJ 25.1436(c)(2)
Pneumatic Systems (Interpretative Material)
See JAR 25X1436(c)(2)
The loads due to vibration and the loads due to temperature effects are those loads which act upon the elements of the system due to environmental conditions.

ACJ 25.1438
Pressurisation and Low Pressure Pneumatic Systems (Acceptable Means of Compliance)
See JAR 25.1438
1 Strength
1.1 Compliance with JAR 25.1309(b) in relation to leakage in ducts and components will be achieved if it is shown that no hazardous effect will result from any single burst or excessive leakage.
1.2 Each element (ducting and components) of a system, the failure of which is likely to endanger the aeroplane or its occupants, should satisfy the most critical conditions of Table 1.
TABLE 1
Conditions 1 Conditions 2

15 P1 at T1 30 P1 at T1
133 P2 at T2 266 P2 at T2 10 P3 at T3 20 P3 at T3
_ 10 P4 at T4
P1 = the most critical value of pressure encountered during normal functioning.
T1 = the combination of internal and external temperatures which can be encountered in association with pressure P1.
P2 = the most critical value of pressure corresponding to a probability of occurrence 'reasonably probable'.
T2 = the combination of internal and external temperatures which can be encountered in association with pressure P2.
P3 = the most critical value of pressure corresponding to a probability of occurrence 'remote'.
T3 = the combination of internal and external temperatures which can be encountered in association with pressure P3.
P4 = the most critical value of pressure corresponding to a probability of occurrence 'extremely remote'.
T4 = the combination of internal and external temperatures which can be encountered in association with pressure P4.
1.3 After being subjected to the conditions given in column 1 of Table 1, and on normal operating conditions being restored, the element should operate normally and there should be no detrimental permanent distortion.
1.4 The element should be capable of withstanding the conditions given in column 2 of Table 1 without bursting or excessive leakage. On normal operating conditions being restored, correct functioning of the element is not required.
1.5 The element should be capable of withstanding, simultaneously with the loads resulting from the temperatures and pressures given in the Table, the loads resulting form -
a. Any distortion between each element of the system and its supporting structures.
b. Environmental conditions such as vibration, acceleration and deformation.
1.6 The system should be designed to have sufficient strength to withstand the handling likely to occur in operation (including maintenance operations).
2 Tests
2.1 Static tests. Each element examined under 1.2 should be static-tested to show that it can withstand the most severe conditions derived from consideration of the temperatures and pressures given in the Table. In addition, when necessary, sub-systems should be tested to the most severe conditions of 1.2 and 1.5. The test facility should be as representative as possible of the aircraft installation in respect of these conditions.
2.2 Endurance tests. When failures can result in hazardous conditions, elements and/or sub-systems should be fatigue-tested under representative operating conditions that simulate complete flights to establish their lives.

ACJ 25.1439(b)(5)
Protective Breathing Equipment (Interpretative Material and Acceptable Means of Compliance)
See JAR 25.1439(b)(5)
1 If a demand system is used, a supply of 300 litres of free oxygen at 70 and 760 mm Hg pressure is considered to be of 15 minutes duration at the prescribed altitude and minute volume. (Interpretative Material.)
2 Any other system such as a continuous flow system is acceptable provided that it does not result in any significant increase in the oxygen content of the local ambient atmosphere above that which would result from the use of a demand oxygen system. (Interpretative Material.)
3 A system with safety over-pressure would be an acceptable means of preventing leakage. (Acceptable Means of Compliance.)
4 A continuous flow system of the closed circuit rebreather type is an acceptable system. (Acceptable Means of Compliance.)

ACJ 25.1441(b)
Oxygen Equipment and Supply (Interpretative Material)
See JAR 25.1441(b)
1 No material should be used which, in direct contact with oxygen, may give off noxious or toxic gases.
2 Couplings and connectors shall be suitable for their intended use and should satisfy the following:
a. Where used in close proximity to one another, they should be protected against incorrect assembly.
b. If intended for connection and disconnection in flight by a crew member they should be -
i Capable of being made and broken with a gloved hand,
ii Automatically ejected, or otherwise provide an obvious indication that the connection is not properly made,
iii Self sealing and designed to minimise leakage during coupling and uncoupling, and
iv So arranged that where electrical connections are also combined the oxygen circuit is complete and sealed before the electrical connections are made.

ACJ 25.1441(c)
Oxygen Equipment and Supply (Interpretative Material)
See JAR 25.1441(c)
1 For the purposes of JAR 25.1439 to 1453 a 'source' may be a single container or a number of containers connected together into a single supply line. A chemical oxygen generator is considered to be a container.
2 In the case of chemically-generated oxygen systems it is not possible to determine directly the quantity available, but for portable units or where oxygen is supplied to more than five occupants from a single supply source, means should be provided readily to determine on the ground and in flight whether the units have been discharged.
3 Indication of the operational state of the flight-crew system and its supplies should be provided at the appropriate flight-crew station.

ACJ 25.1441(d)
Oxygen Equipment and Supply (Interpretative Material)
See JAR 25.1441(d)
In assessing the required oxygen flow rates and equipment performance standards, consideration should be given to the most critical cabin altitude/time-history following any failure, not shown to be Extremely Improbable, which will result in the loss of cabin pressure taking into account the associated emergency procedures.

ACJ 25.1443
Minimum Mass Flow of Supplemental Oxygen (Interpretative Material)
See JAR 25.1443
The curves shown in Figure 3 may be used to determine the minimum oxygen flow rates required.
These curves represent the theoretical minimum flow
rates, and should therefore be factored to suit the
efficiency of the particular equipment considered.
LEGEND
___ . ___ . ___ . 1443 (a)
____________ 1443 (b)
_ _ _ _ _ _ _ _ 1443(c) 1 and 2

FIGURE 3

ACJ 25.1445(a)(2)
Equipment Standards for the Oxygen Distributing System (Interpretative Material)
See JAR 25.1445(a)(2)
Where the separate reserve is provided by an automatic device, means should be provided to override this device manually unless its failure can be shown to be Extremely Remote.

ACJ 25.1445(a)(3)
Equipment Standards for the Oxygen Distributing System (Interpretative Material)
See JAR 25.1445(a)(3)
The requirements of this paragraph need not be applied to chemical generator systems which supply not more than five occupants from a single supply source.

ACJ 25.1447(c)
Equipment Standards for Oxygen Dispensing Units (Interpretative Material)
See JAR 25.1447(c)
Where National Operational Regulations do not require all passengers to be provided with oxygen, (c)(3) and (c)(4) may not apply.

ACJ 25.1447(c)(1)
Equipment Standards for Oxygen Dispensing Units (Interpretative Material)
See JAR 25.1447(c)(1)
1 When oxygen masks are presented, oxygen should be supplied to the mask but without flow.
2 Oxygen flow from the mask should be initiated automatically on pulling the mask to the face.
3 Facilities for manual presentation by a crew member should be provided on each dispensing unit.
4 Indication of the operation of the automatic presentation system should be provided at the appropriate flight-crew station.
5 The design of the automatic presentation system should take into account that when the landing field altitude is less than 2000 feet below the normal preset automatic presentation altitude, the automatic presentation altitude may be reset to landing field altitude plus 2000 feet.

ACJ 25.1447(c)(2)
Equipment Standards for Oxygen Dispensing Units (Interpretative Material)
See JAR 25.1447(c)(2)
Unless it is required that the pilot at the control is wearing his mask and breathing oxygen while the altitude exceeds 25 000 feet, the design of the flight-crew masks and their stowages should be such that each mask can be placed in position and put into operation in not more than five seconds, one hand only being used, and will thereafter remain in position, both hands being free.

ACJ 25.1447(c)(4)
Equipment Standards for Oxygen Dispensing Units (Interpretative Material)
See JAR 25.1447(c)(4)
1 The equipment should be so located as to be within reach of the cabin attendants while seated and restrained at their seat stations.
2 The mask/hose assembly should be already connected to the supply source, and oxygen should be delivered with no action being required except turning it on and donning the mask.
3 Where a cabin attendant's work area is not within easy reach of the equipment provided at his seat station, an additional unit should be provided at the work area.

ACJ 25.1449
Means for Determining Use of Oxygen (Interpretative Material)
See JAR 25.1449
A simple flow indicator should be provided unless it can be shown that the inflation of the economiser system provides effective indication. A system using a simple re-breathing bag would not be acceptable.

ACJ 25.1450
Chemical Oxygen Generators (Interpretative Material)
See JAR 25.1450
Where sustained operation is achieved by successive replacement of generator elements, adequate provision should be made for the handling and disposal of used elements.

ACJ 25.1453
Protection of Oxygen Equipment from Rupture (Interpretative Material)
See JAR 25.1453
1 Parts of the system subjected to high oxygen pressure should be kept to a minimum and should be remote from occupied compartments. Where such parts are installed within occupied compartments they should be adequately protected from accidental damage.
2 Each container, component, pipe and coupling should have sufficient strength to withstand a pressure equivalent to not less than the maximum working pressure acting on that part of the system when multiplied by the appropriate Proof and Ultimate factors given in Table 1. The maximum working pressure includes tolerances of any pressure limiting means and possible pressure variations in the normal operating modes. Account should also be taken of the effects of temperature up to the maximum anticipated temperature to which the system may be subjected. Transient or surge pressures need not be considered except where these exceed the maximum working pressure multiplied by 110.
TABLE 1
Systems Element Proof Factor Ultimate Factor
Containers 15 20
Flexible hoses 20 40
Pipes and couplings 15 30
Other components 15 20

3 Each source should be provided with a protective device (e.g. rupture disc). Such devices should prevent the pressure from exceeding the maximum working pressure multiplied by 15.
4 Pressure limiting devices (e.g. relief valves), provided to protect parts of the system from excessive pressure, should prevent the pressures from exceeding the applicable maximum working pressure multiplied by 133 in the event of malfunction of the normal pressure controlling means (e.g. pressure reducing valve).
5 The discharge from each protective device and pressure limiting device should be vented overboard in such a manner as to preclude blockage by ice or contamination, unless it can be shown that no hazard exists by its discharge within the compartment in which it is installed. In assessing whether such hazard exists consideration should be given to the quantity and discharge rate of the oxygen released, the volume of the compartment into which it is discharging, the rate of ventilation within the compartment and the fire risk due to the installation of any potentially flammable fluid systems within the compartment.
6 In addition to meeting the requirements of JAR 25.1453, oxygen containers may have to be approved in accordance with national regulations.
NOTES:
1 The proof pressure should not cause any leakage or permanent distortion.
2 The ultimate pressure should not cause rupture but may entail some distortion.
[ ACJ 25.1457
Cockpit Voice Recorders (Interpretative Material)
See JAR 25.1457
In showing compliance with JAR 25.1457, the applicant should take account of EUROCAE document No. ED-56 'Minimum Operational Performance Requirement for Cockpit Voice Recorder System', which will be referred to in a JAR TSO when published. ]
[ ACJ 25.1459(a)(4)
Flight Recorders (Acceptable Means of Compliance)
See JAR 25.1459(a)(4)
An acceptable means of compliance would be to provide a combination of system monitors and built-in test functions which would detect and indicate the following:
a. Loss of electrical power to the flight recorder system.
b. Failure of the data acquisition and processing stages.
c. Failure of the recording medium and/or drive mechanism.
d. Failure of the recorder to store the data in the recording medium as shown by checks of the recorded data including, as reasonably practicable for the storage medium concerned, correct correspondence with input data. ]

ACJ 25.1459(b)
Flight Recorders (Acceptable Means of Compliance)
See JAR 25.1459(b)
1 The phrase 'as far aft as practicable' should be interpreted as a position sufficiently aft as to be consistent with reasonable maintenance access and in a position to minimise the probability of damage from crash impact and subsequent fire.
2 The container should remain attached to the local structure under normal, longitudinal and transverse accelerations of at least 10 g.

ACJ 25.1499(a)
Domestic Services and Appliances (Interpretative Material)
See JAR 25X1499(a)
When considering failures of domestic appliances, the effect of the loss of the water supply to a water heater, with the electrical supply maintained, should be taken into account.

ACJ 25.1499(b)
Domestic Services and Appliances (Acceptable Means of Compliance)
See JAR 25X1499(b)
The design of galley and cooking appliance installations should be such as to facilitate cleaning to limit the accumulation of extraneous substances which may constitute a fire risk.
ACJ - Subpart G

ACJ 25.1501
Operating Limitations and Information - General (Interpretative Material)
See JAR 25.1501
The limitations and information established in accordance with Subpart G should be only those which are within the competence of the flight crew to observe, and should relate only to those situations (including pre- and post-flight) with which a flight crew member might reasonably be concerned.

ACJ 25.1516
Other Speed Limitations (Interpretative Material)
See JAR 25X1516
Speed limitations for devices such as spoilers, speed brakes, high lift devices, thrust reversers, landing lights and the opening of doors and direct vision windows, should be included.

ACJ 25.1517
Rough Air Speed (Interpretative Material)
See JAR 25X1517
VRA should be less than VMO - 35 knots TAS.

ACJ 25.1519
Weight, Centre of Gravity and Weight Distribution (Interpretative Material)
See JAR 25.1519
A statement of the maximum certificated take-off and landing weights, and the minimum certificated take-off and landing weights, should be established, together with the maximum ramp or taxying weight, the maximum zero-fuel weight and any other fixed limit on weight, including weight limitations resulting from such factors as brake energy limits, tyre limits, etc., established in accordance with the airworthiness standards of JAR-25. Any limitations on aeroplane loading associated with the stated weight limitations (e.g. fuel load and usage, maximum fuel for landing) should be considered.

ACJ 25.1521
Power-Plant Limitations (Interpretative Material)
See JAR 25.1521
1 In furnishing limitations, consideration should be given to the following. The list does not necessarily include all the items to be considered for a given aeroplane.
a. Rotational speeds.
b. Exhaust and/or turbine gas temperature.
c. Oil temperatures and pressures.
d. Fuel temperatures and pressures.
e. Water and/or water methanol usage.
f. Anti-icing.
g. Specifications of approved fuels, oils and additives.
2 Other parameters, e.g. time, altitude, ambient temperatures, airspeed, may be necessary in defining power-plant limitations.
3 All operating phases should be considered in establishing the power-plant limitations.

ACJ 25.1523
Minimum Flight Crew (Interpretative Material)
See JAR 25.1523
1 Both the number and identity of the flight crew members should be established.
2 If the minimum flight crew varies with the kinds of operation to which the aeroplane is limited, the approved number and identity of the flight crew members should be stated for each kind of operation.
3 If a particular flight crew member's station has to be occupied at all material times, this should be stated when specifying the minimum flight crew.

ACJ 25.1524
Systems and Equipment Limitations (Interpretative Material)
See JAR 25X1524
Examples of systems and equipment installations for which limitations may be appropriate include, but are not limited to, electrical, fuel, hydraulic, pneumatic, cabin pressurisation, air conditioning, airframe fire protection, airframe ice protection, anti-skid and auto-braking systems, also autopilot, auto-throttle, flight director, approach coupler and yaw damper.

ACJ 25.1541
Markings and Placards - General (Interpretative Material)
See JAR 25.1541
Markings or placards should be placed close to or on (as appropriate) the instrument or control with which they are associated. The terminology and units used should be consistent with those used in the Flight Manual. The units used for markings and placards should be those that are read on the relevant associated instrument.

ACJ 25.1543
Instrument Markings - General (Interpretative Material)
See JAR 25.1543
The markings should be such that the instrument remains easily readable with the minimum of confusion.

ACJ 25.1545
Airspeed Limitation Information (Interpretative Material)
See JAR 25.1545
A placard could be used when the speed limitation can be a simple presentation (e.g. an IAS speed up to a given altitude and an indicated Mach number thereafter). A complex speed limitation should be presented automatically on the instrument, (e.g. by means of an additional moving pointer).

ACJ 25.1549
Powerplant Instruments (Interpretative Material)
See JAR 25.1549
1 Powerplant instrument range markings are intended to indicate to flight crew members, at a glance, that the powerplant operation is being accomplished in a safe or desirable, undesirable but allowable, or unsafe region. The colour red indicates an unsafe condition which requires immediate and precise action by the flight crew. The use of multiple red lines should be avoided to minimise confusion.
2 A precautionary range is a range where limited operation is permissible, as indicated in the aeroplane Flight Manual. Experience has shown that to satisfy the requirement for clearly visible markings, the following minimum dimensions should be observed.
[ a. Red, yellow and green lines. 005 in wide and 03 in long. ]
b. Red, yellow and green arcs and areas. 01 in wide, length as required.

ACJ 25.1557(a)
Baggage and Cargo Compartment and Ballast Location (Acceptable Means of Compliance)
See JAR 25.1557(a)
If baggage, cargo compartment and ballast location limitations are complex and involve, for example, additional limitations on loading intensity and distribution, it is acceptable to provide a placard making reference to the appropriate document.

ACJ 25.1581(a)
Flight Manual - General (Acceptable Means of Compliance)
See JAR 25.1581(a)
The layout of the Flight Manual should be in accordance with the Provisional Acceptable Means of Compliance for the standardisation of approved aeroplane Flight Manuals, as prescribed in ICAO Airworthiness Technical Manual 9051 - AN 896. The terminology and units used in the Flight Manual should be in accordance with the recommendation of ICAO Airworthiness Technical Manual 9051 - AN 896 or as prescribed by the Authority.

ACJ 25.1583(c)
Centre-of-Gravity Limitations (Interpretative Material)
See JAR 25.1583(c)
1 Indication should be given in tabular or graphic form of the c.g. limits for take-off and landing and for any other practicably separable flight condition, as appropriate for the range of weights between the maximum take-off weight and the minimum landing weight presented in accordance with JAR 25.1583(c). The landing gear position appropriate to each condition should be shown, or, alternatively, data should be presented for landing-gear-extended position only and should include the moment change due to gear retraction. C.g. limits should be presented in terms of both distance-from-datum and percentage of the mean aerodynamic chord (MAC). The datum for the former should be defined and the length and location of the MAC should be stated.
2 For those weight limitations which vary with runway length, altitude, temperature and other variables the variation in weight limitation may be presented as graphs in the performance section of the Flight Manual, and included as limitations by specific reference, in the limitations section, to the appropriate graph or page.

ACJ 25.1583(i)
Manoeuvring Flight Load Factors (Interpretative Material)
See JAR 25.1583(i)
The flight manoeuvring limit load factors for which the structure is approved, expressed in terms of normal acceleration, or g, should be included. If more restrictive flight load factors are established for particular operations outside the normal operating envelope (e.g. landing flap position with maximum take-off weight) such factors should be presented and defined.
[ ACJ 25.1583(k)
Maximum Depth of Runway Contaminants for Take-off Operations (Acceptable Means of Compliance)
See JAR 25.1583(k)
Compliance with JAR 25.1583(k) may be shown using either Method 1 or Method 2 -
a. Method 1. If information on the effect of runway contaminants on the expected take-off performance of the aeroplane is furnished in accordance with the provisions of JAR 25X1591(c)(2), take-off operation should be limited to the contamination depths for which take-off information is provided.
b. Method 2. If information on the effect of runway contaminants on the expected take-off performance of the aeroplane in accordance with the provisions of JAR 25X1591(c)(2) is not provided, take-off operation should be limited to runways where the degree of contamination does not exceed the equivalent of 3 mm (0125 inch) of water, except in isolated areas not exceeding a total of 25% of the area within the required length and width being used.
NOTE 1 In establishing the maximum depth of runway contaminants it may be necessary to take account of the maximum depth for which the engine air intakes have been shown to be free of ingesting hazardous quantities of water or other contaminants in accordance with JAR 25.1091(d)(2).
NOTE 2: Unless performance effects are based on tests in water depths exceeding 15mm, or on other evidence equivalent in accuracy to the results of direct testing, it will not normally be acceptable to approve take-off operation in depths of contaminants exceeding the equivalent of 15mm of water. ]

ACJ 25.1585(a)
Operating Procedures (Interpretative Material)
See JAR 25.1585(a)
1 In furnishing information and instructions, consideration should be given to the following. The lists do not necessarily include all items to be considered for a given aeroplane. The categorisation of certain items may need to be modified because of design features or other considerations.
2 Emergency Procedures
a. Engine and APU fire/separation/severe damage
b. Smoke or fire in cockpit/cabin/cargo compartment
c. Rapid decompression/emergency descent
d. Landing or go-around with jammed stabiliser
e. Runaway stabiliser
f. Flight with all engines inoperative
g. Ditching
3 Other Procedures
a. Engine starting
b. APU operation
c. Fuel management. The effect on unusable fuel quantity due to fuel booster pump failure should be stated.
d. Reverse thrust system
e. Navigation system
f. Rain repellent system
g. Automatic flight control systems
h. Cabin pressurisation system
i. Oxygen system
j. Hydraulic system
k. Electrical system
l. Anti-ice/de-ice system
m. Operation in turbulence
n. Equipment cooling
o. Flight controls
p. Stall warning/stall identification system
q. Braking system
r. Fuel dumping
s. Go-around with minimum fuel
t. Landing in abnormal configurations
u. Engine shut-down and relight in flight
v. Approach and landing with engine(s) inoperative
w. Go-around with engine(s) inoperative
x. Landing gear alternate operation
4 Certain items listed in 3 may also need to be considered under 2.
5 Observance of these procedures may not be mandatory and approval of such procedures is not intended to prohibit or discourage development and use of improved or equivalent alternative procedures based on operational experience with the aeroplane.
6 The procedures to be followed by the flight crew in the event of an engine fire, severe damage or separation of the engine should be similar, and should include identification of the failed engine as the primary action as far as the powerplant is concerned.

ACJ 25.1585(c)
Cruise Manoeuvring Capability (Acceptable Means of Compliance)
See JAR 25.1585(c)
The buffet onset envelopes should be accompanied by information on the maximum altitude at which it is possible to achieve a positive normal acceleration increment of 03 g without exceeding the buffet onset boundary, at any given combination of weight, centre of gravity location and airspeed. (See also ACJ 25.251(e).)

ACJ 25.1587(b)(1)
Climb one-engine-inoperative (Acceptable Means of Compliance)
See JAR 25.1587(b)(1)
The bank angle used in showing compliance with JAR 25.121 should be scheduled in the Flight Manual. Where it is more practical to quote the degree of lateral control (e.g. control wheel level) instead of the bank angle, this would be acceptable.
[ ACJ 25.1587(b)(7)
Performance Information: Presentation of Landing Distance (Acceptable Means of Compliance)
See JAR 25.1587(b)(7)
1 The landing distance from a height of 50 ft determined in accordance with JAR 25.125 should be presented together with associated conditions for weights up to the maximum take-off weight, standard temperature and corrected for not more than 50% of nominal headwind component, and not less than 150% of nominal tailwind component.
2 Data should be presented for level, smooth, dry, hard-surfaced runways. At the option of the applicant, additional data may be presented to show the effect of runway slope and temperature, within the operational limits of the aeroplane.
3 To facilitate application of operating regulations, the landing distance may be presented in the form of the operational or "factored" runway length, using the appropriate factors prescribed by the operating regulations of the state of the registry of the aeroplane. The factors applied should be stated together with associated conditions. ]
[ ACJ - SUBPART J

ACJ 25.901(b)(2)
Assembly of Components (Auxiliary Power Units) (Interpretative Material)
See JAR 25A901(b)(2)
The objectives of JAR 25.671(b) should be satisfied with respect to APU systems, where the safety of the aeroplane could otherwise be jeopardised.

ACJ 25.901(b)(4)
Electrical Bonding (Auxiliary Power Units) (Interpretative Material)
See JAR 25A901(b)(4)
Where the APU is not in direct electrical contact with its mounting the engine should be electrically connected to the main earth system by at least two removable primary conductors, one on each side of the APU.

ACJ 25.901(d)
General (Auxiliary Power Units) (Interpretative Material)
See JAR 25A901(d)
The need for additional tests, if any, in hot climatic conditions should take account of any tests made by the APU constructor to establish APU performance and functioning characteristics and of satisfactory operating experience of similar power units installed in other types of aeroplane.
The applicant should declare the maximum climatic conditions for which compliance will be established and this should not be less severe than the ICAO Intercontinental Maximum Standard Climate (100F (378C) at sea-level). If the tests are conducted under conditions which deviate from the maximum declared ambient temperature, the maximum temperature deviation should not normally exceed 25F (1388C).

ACJ 25.903(e)(2)
APUs (Auxiliary Power Units) (Interpretative Material)
See JAR 25B903(e)(2)
1 General. The minimum acceptable relight envelope is defined in paragraph 2.
2 Envelope of Altitude and Airspeed
2.1 Sufficient flight tests should be made over the range of conditions detailed in 2.2 and 2.3 to establish the envelope of altitude and airspeed for reliable APU restarts, taking into account the results of restart tests completed by the APU constructor on the same type of APU in an altitude test facility or flying test bed, if available, and the experience accumulated in other aircraft with the same APU. The effect of APU deterioration in service should be taken into account.
2.2 Altitude and Configuration. From sea-level to the maximum declared restarting altitude in all appropriate configurations likely to affect restarting, including the emergency descent configuration.
2.3 Airspeed. From the minimum to the maximum declared airspeed at all altitudes up to the maximum declared APU restarting altitude. The airspeed range of the declared relight envelope should cover at least 30 kt. ]
[ 2.4 Delay Tests. The tests referred to in 2.2 should include the effect on APU restarting performance of a delay period between APU shut-down and restarting.

ACJ 25.939(a)
Turbine APU Operating Characteristics (Auxiliary Power Units) (Interpretative Material)
See JAR 25A939(a)
The wording 'in flight' should be interpreted to cover all operating conditions from APU start until shut-down.

ACJ 25.943
APU Operating Characteristics (Auxiliary Power Units) (Interpretative Material)
See JAR 25A943
1 Compliance with JAR 25A943 should be shown by design analysis and flight tests. The flight tests should include manoeuvre in which less than zero 'g' occurs for one continuous period of at least 5 seconds and a further manoeuvre with two periods of less than zero 'g' with a total time for these two periods of at least 5 seconds.
2 In the case of non-essential APUs, inadvertent shut-down due to negative accelerations is acceptable.

ACJ 25.953(b)
Fuel System Independence (Auxiliary Power Units) (Interpretative Material)
See JAR 25A953(b)
The fuel supply to an APU may be taken from the fuel supply to the main engine if provision is made for a shut-off means to isolate the APU fuel line.

ACJ 25.961(a)(5)
Fuel System Hot Weather Operation (Auxiliary Power Units)(Acceptable Means of Compliance)
See JAR 25B961(a)(5)
Subject to agreement with the Authority, fuel with a higher vapour pressure may be used at a correspondingly lower fuel temperature provided the test conditions closely simulate flight conditions corresponding to an initial fuel temperature of 110F (433C) at sea-level.

ACJ 25.991
Fuel Pumps (Auxiliary Power Units)(Interpretative Material)
See JAR 25B991
If the fuel supply to the APU is taken from the fuel supply to the main engine, no separate pumps need be provided for the APU. ]
[ ACJ 25B1093(b)(2)
APU Air Intakes (Auxiliary Power Units)(Acceptable Means of Compliance and Interpretative Material)
See JAR 25B1093(b)(2)
1 General. Two ways of showing compliance with JAR 25B1093(b)(2) are given.
1.1 Method 1. Method 1 is an arbitrary empirical method based on United Kingdom and French practice. This method is acceptable to all participating countries.
1.2 Method 2. Method 2 is a general approach based on US practice in applying FAR Part 25, Appendix C. If this method is used, each application will have to be evaluated on its merits.
2 Method 1 (Acceptable Means of Compliance)
2.1 In establishing compliance with the requirements of JAR 25B1093(b)(2), reference should be made to ACJ 25.1419, paragraph 1.
2.2 The intake may be tested with the APU in accordance with the requirements of JAR-APU, Section 1, paragraph 5.2 and the Advisory Material for the testing of APUs in Icing Conditions.
2.3 When the intake is assessed separately it should be shown that the effects of intake icing would not invalidate the icing tests of JAR-APU. Factors to be considered in such evaluation are -
a. Distortion of the airflow and partial blockage of the intake.
b. The shedding into the APU of intake ice of a size greater than the APU is known to be able to ingest.
c. The icing of any APU sensing devices, other subsidiary intakes or equipment contained within the intake.
d. The time required to bring the protective system into full operation.
2.4 Tests in Ice-forming Conditions. An acceptable method of showing compliance with the requirements of JAR 25B1093(b)(2), including Appendix C, is given in this paragraph.
2.4.1 When the tests are conducted in non-altitude conditions, the system power supply and the external aero-dynamic and atmospheric conditions should be so modified as to represent the required altitude conditions as closely as possible. The altitudes to be represented should be as indicated in Table 1 for simulated tests, or that appropriate to the desired temperature in flight tests, except that the test altitude need not exceed any limitations proposed for approval. The appropriate intake incidences or the most critical incidence, should be simulated.
2.4.2 Two tests (which may be separated or combined) should be conducted at each temperature condition of Table 1, at or near the indicated altitude -
a. 30 minutes in the conditions of Table 1 column (a) appropriate to the temperature.
b. Three repetitions of 5 km in the conditions of Table 1, column (b), appropriate to the temperature followed by 5 km in clear air. ]
TABLE 1
____________________________________________________________________
Altitude Liquid water
Ambient air content (g/m3) Mean effective
temperature droplet diameter
(C) (ft) (m) (a) (b) (m)
-10 17 000 5200 0.6 2.2
-20 20 000 6100 0.3 1.7 20
-30 25 000 7600 0.2 1.0

2.4.3 At the conclusion of each of the tests of 2.4.2 the ice accretion should be such as not to adversely affect the subsequent running and functioning of the APU.
2.4.4 If the APU intake contains features or devices which could be affected by freezing fog conditions then in addition to the above tests of 2.4.2 a separate test on these parts or devices should be conducted for a duration of 30 minutes with the heat supply to the tested parts as would be available with the APU set to the minimum ground idle conditions approved for use in icing in an atmosphere of -2C and a liquid water concentration of 03 g/m3. The mean effective droplet size for the test should be 20m. At the end of the period the ice accretion on the tested part should not prevent its proper functioning nor should the ice be of such size as to hazard the APU if shed.
3 Method 2 (Interpretative Material)
3.1 In establishing compliance with the requirements of JAR 25B1093(b)(2), reference should be made to JAR 25.1419 and ACJ 25.1419.
3.2 The intake may be tested with the APU in accordance with a programme of tests which results from an analysis of the icing conditions and the APU conditions appropriate to the installation.
3.3 When the intake is assessed separately it should be shown that the effects of intake icing would not invalidate any APU certification tests. Factors to be considered in such evaluation are -
a. Distortion of the airflow and partial blockage of the intake.
b. The shedding into the APU of intake ice of a size greater than the APU is known to be able to ingest.
c. The icing of any APU sensing devices, other subsidiary intakes or equipment contained within the intake.
d. The time required to bring the protective system into full operation.
3.4 When tests are conducted in non-altitude conditions, the system power supply and the external aerodynamic and atmospheric conditions should be so modified as to represent the altitude condition as closely as possible. The appropriate intake incidences or the most critical incidence, should be simulated.
3.5 Following the analysis required in JAR 25.1419(b), which will determine the critical icing conditions within the envelope of icing conditions defined by Appendix C Figures 1 to 3 and Appendix C Figures 4 to 6, tests should be conducted at such conditions as are required to demonstrate the adequacy of the design points.
3.6 At the conclusion of each of the tests the ice accretion should be such as not to adversely affect the subsequent running and functioning of the APU.
3.7 If the APU intake contains features or devices which could be affected by freezing fog conditions then a separate assessment for these parts should be conducted assuming a duration of 30 minutes and an atmosphere of -2C and a liquid water concentration of 03 g/m3, with the heat supply to the part as would be available with the APU set to the minimum ground idle conditions approved for use in icing. The mean effective droplet size should be 20m. At the end of the period the ice accretion on the part should not prevent its proper functioning, nor should the ice be of such size as to hazard the engine if shed.

ACJ 25.1195(b)
Fire Extinguisher Systems (Auxiliary Power Units)(Interpretative Material and Acceptable Means of Compliance)
See JAR 25A1195(b)
Acceptable methods to establish the adequacy of the fire extinguisher system are laid down in Advisory Circular 20 - 100.
ACJ - Appendices
ACJ to Appendix F, Part IV
Test Method to Determine the Heat Release Rate from Cabin Materials Exposed to Radiant Heat. (Acceptable Means of Compliance)
See Appendix F, Part IV
Appendix F, Part IV (b)(4) Air Distribution System.
The air distribution is to be determined by the equipment design. The 3-to-1 ratio described in this section is approximate. An external air distribution system which will deliver that ratio precisely is not permitted as a substitute for the air distributor plates.
Appendix F, Part IV (b)(6) Specimen Holders.
In order to accommodate specimens which distort and delaminate during testing, two 0020-inch (0508mm) stainless steel wires should be used to secure the specimens to the holder during the testing.
These wires should be used with all specimens and are in addition to the drip pan that should be used for materials which are prone to melting and dripping.
Appendix F, Part IV (b)(8) Pilot-Flame Positions.
Various installations have experienced difficulties with the pilot burners being extinguished during the test.
The following revisions to the pilot burner configurations have been found to be acceptable:
(1) For the lower pilot burner - a sparking device which either sparks automatically at approximately to 1 second intervals or is manually operated, which requires continuous monitoring of the pilot flame.
Note: This requires that the laboratory test procedure specifies that the technician must continuously monitor the pilot for each test and that failure to do so will invalidate the test results.
(2) For the upper pilot burner - a manual or automatic sparking device or a revision to the hole system in the burner. One approved deviation utilises 14 holes using a number 59 drill bit.
Appendix F, Part IV (c)(1) Heat Release Rate.
The use of a flowmeter is not acceptable.
The thermopile voltage should be measured for 10 seconds and then averaged.
Appendix F, Part IV (e) Procedure.
The outer door should be closed between tests to maintain the heat within the chamber. It is recommended that the outer door be hinged to facilitate implementing this recommendation. If a detachable door is used, a separate door should be installed during sample holder preparation and installation. This recommendation is based on the 40-seconds holding time (60 seconds less 20 seconds of data acquisition time) required in (e)(4), being insufficient to allow the chamber to reach equilibrium, if the outer door is open for too long between tests.
Appendix F, Part IV (f) Calculations.
It has been found that a typical range for the calibration factor is 8 to 15. If a calibration factor is calculated which falls outside this range, the calculation should be reviewed.
If the factor continues to fall outside this range, the appropriate Authority should be contacted.

LAST UPDATE:  [an error occurred while processing this directive]
AUTHOR:  Prof. Dr. Scholz
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home  Prof. Dr. Scholz
home  Aircraft Design and Systems Group (AERO)
home  Aeronautical Engineering   deutsch
home  Department of Automotive and Aeronautical Engineering  deutsch
home  Faculty of Engineering and Computer Science
home  Hamburg University of Applied Sciences