150 criteria depicted in figure 31-2 for in-trim maneuver stability. Slight control force reversals, as discussed in paragraph (a), above, will be permitted for speeds between VFC/MFC and VDF/MDF only if:
1 No severe longitudinal control force slope reversals exist;
2 Any pitching tendency (uncommanded changes in load factor) should be mild and readily controllable; and
3 The airplane’s pitch response to primary longitudinal control should be predictable to the pilot.
Figure 33-1. Mistrimmed Maneuvering Characteristics
Speeds Between VFC/MFC and VDF/MDF
(3) Section 25.255(c) requires that the investigation of maneuvering stability (§ 25.255(b)) include all attainable acceleration values between -lg and +2.5g. Sections
25.333(b) and 25.337, to which it refers, limit the negative g maximum to 0g at VD. Section
25.25l further limits the g to that occurring in probable inadvertent excursions beyond the buffet onset boundary at those altitudes where buffet is a factor.
(4) Section 25.255(c)(2) allows for extrapolation of flight test data by an acceptable method. For example, if the stick force gradient between 0 and +2g agrees with predicted data, extrapolation to -1g and 2.5g should be allowed.
(5) Section 25.255(d) requires flight tests to be accomplished from the normal acceleration at which any marginal stick force reversal conditions are found to exist to the applicable limits of § 25.255(b)(1). This requirement takes precedence over the extrapolation allowance described in paragraph (4), above.
(6) Section 25.255(e) , limits the investigation to the required structural strength limits of the airplane and maneuvering load factors associated with probable inadvertent excursions beyond the boundary of the buffet onset envelope. It also accounts for the fact that speed may increase substantially during test conditions in the -1g to +1g range. It limits the entry speed to avoid exceeding VDF/MDF.
(7) Section 25.255(f) requires that in the out-of-trim condition specified in § 25.255(a), it must be possible to produce at least 1.5g during recovery from the overspeed condition of VDF/MDF. If adverse flight characteristics preclude the attainment of this load factor at the highest altitude reasonably expected for recovery to be initiated at VDF/MDF following an upset at high altitude, the flight envelope (c.g., VDF/MDF, altitude, etc.) of the airplane should be restricted to a value where 1.5g is attainable. If trim must be used for the purpose of obtaining
1.5g, it must be shown to operate with the primary control surface loaded to the least of three specified values.
(a) The force resulting from application of the pilot limit loads of § 25.397 (300 lbs.).
(b) The control force required to produce 1.5g (between 125 and 300 lbs.).
(c) The control force corresponding to buffeting or other phenomena of such intensity that it is a strong deterrent to further application of primary longitudinal control force.
d. Procedures .
(1) Basic compliance is determined by the characteristics of FS/g (normally a plot).
Any standard flight test procedure that yields an accurate evaluation of FS/g data in the specified range of speeds and acceleration should be considered for acceptance. Bounds of investigation and acceptability are set forth in the rule and in discussion material above, and broad pilot discretion is allowed in the selection of maneuvers.
(2) Investigation Range . Out-of-trim testing should be done at the most adverse loading for both high and low control forces. Testing should be accomplished both at the “Q” and Mach limits.
(3) The ability to actuate the primary controls (including trim), when loaded, should be considered prior to the tests.
Chapter 3 - Structure
[Reserved]
Chapter 4 – Design and Construction
Section 1. General [Reserved]
Section 2. Control Surfaces [Reserved]
Section 3. Control Systems
34. General - § 25.671.
a. Explanation . This material deals with the all-engines out case of § 25.671(d). The intent of this rule is to assure that in the event of failure of all engines, the airplane will be controllable and an approach and landing flare possible. This may be done by analysis where the method is considered reliable.
b. Procedures . The airplane should be evaluated to determine that:
(1) It is controllable following the failure of all engines in the climb, cruise, descent, approach, and holding configurations and can be flared to a landing attitude from a reasonable approach speed. The airplane must be controllable when all engines fail in each of the specified configurations and in any specific configuration that is to be selected and maintained following the failure of the engines in accordance with the AFM emergency operating procedures.
(2) The effectiveness of the emergency power to drive the airplane control system, whether generated from a windmilling engine or an auxiliary power supply, should be demonstrated in flight.
(3) For airplanes with fully powered or electronic flight control systems, the emergency procedures section of the approved AFM should contain the appropriate operating procedures and a statement similar to the following:
“The airplane has a fully powered (or electronic) control system that is dependent upon engine windmill RPM, or an auxiliary power supply, to provide the necessary source of control system power in the event all engines fail in flight. A minimum airspeed of XXX knots IAS will provide adequate hydraulic or electrical power for airplane controllability in this emergency condition.”
35. - 45. [Reserved]
46. Flap and Slat Interconnections - § 25.701.
a. Explanation . In accordance with § 25.701(a), if the wing flaps are not mechanically interconnected, tests or analysis should be conducted to simulate appropriate flap malfunctioning
154 during takeoffs, approaches, and landings to demonstrate that the airplane is safe under these conditions. This demonstration should not require exceptional pilot skills or strength.
b. Reference . See paragraph 20a(3) of this AC and AC 25-14, “High Lift and Drag Devices,” dated May 4, 1988, for additional guidance.
c. Procedures . As appropriate to the installation.
47. Takeoff Warning System - § 25.703. [Reserved]
Section 4. Landing Gear
48. - 51. [Reserved]
52. Retracting Mechanism - § 25.729.
a. Explanation . None.
b. Procedures .
(1) In accordance with the provisions of § 25.729, flight tests should be conducted to demonstrate the ability of the landing gear and associated components, in their heaviest configuration, to properly extend and retract at:
(a) V LO (the placard airspeed) in the cruise configuration at near 1g flight and normal yaw angles; and
(b) airspeeds and flap settings corresponding to typical landings. The landing gear operating placard airspeed, VLO, established in accordance with § 25.1515(a), should not exceed the 1.6 VS1 design value of § 25.729(a)(1)(ii).
NOTE: “Normal” yaw angles are those a ssociated with engine-out flight and counteracting crosswinds of up to 20 knots.
(2) The alternate extend system should be demonstrated at airspeeds up to V LO at near 1g flight and normal yaw angles. An envelope of emergency extension capability should be established and presented in the emergency operating procedures section of the AFM. (Refer to NOTE paragraph above.)
(3) Operation test.- § 25.729(d).
(a) The engine-out gear retraction time should be determined from flight tests with one engine at idle power and the operating engine(s) adjusted to provide the lowest thrustto-weight ratio to be certificated. The hydraulic system should be in the critical configuration corresponding to an actual engine failure condition. The airplane should be stabilized on a steady heading before gear retraction is initiated. The resulting gear retraction time will be used in developing AFM takeoff flight path performance information in accordance with § 25.111.
(b) Gear retraction time is the time from landing gear lever movement to the “UP” position until the last landing gear, including doors, is in the retracted configuration. Allowance should be made for any delays associated with the landing gear indication system.
(4) Position indication and aural warning, § 25.729(e): It should be confirmed that the actual landing gear position agrees with the position indicated on the landing gear indicator.
Landing gear aural warning should meet the intent of § 25.729(e)(2) through (e)(4). A combination of flight tests, ground tests, and analysis may be used to show compliance with these requirements.
53. Wheels - § 25.731.
a. References .
(1) TSO-C135a “Transport Airplane Wheels and Wheel and Brake Assemblies,” dated July 1, 2009.
(2) Paragraph 55b(4)(f) of this AC, Wheel Fuse Plugs.
(3) Paragraph 55c(7) of this AC, Wheel Fuse Plug Design.
(4) AC 21-29C, Change 1, “Detecting and Reporting Suspected Unapproved Parts,” dated July 22, 2008.
b. Explanation .
(1) Background .
(a) Original guidelines for wheels in § 4b.335 of the Civil Air Regulations (CAR) were superseded by TSO-C26 and subsequent revisions. Early versions of TSO-C26 referred to minimum standard requirements in versions of Society of Automotive Engineers (SAE) Aeronautical Standard (AS) 227. Minimum standards were subsequently specified in TSO-C26b and later revisions. For braked wheels, wheels and brakes must be approved under § 25.735(a) as an assembly in recognition of design and safety interdependencies associated with thermal control, vibration control, structural stresses, etc. References (2) and (3) of paragraph a, above, provide insight into the criticality of proper fuse function to support airworthiness. As demands
156 increased for longer life wheels and more robust designs, changes were introduced such as the addition of combined vertical and side loads for a portion of roll testing; an increase in roll miles at maximum static load from 1,000 to 2,000 miles; and addition of a roll-on-rim requirement to increase robustness in wheel flange areas.
(b) In most cases, wheels are usually removed from service based upon condition.
Typically, inspection frequency will increase as life on a particular wheel is accumulated in accordance with the wheel and brake suppliers component maintenance manual (CMM). If wheel failures occur on the airplane, they typically occur when the airplane is on the ground with wheel and tire assembly loaded.
(c) The trend in certain airplane manufacturer specifications is to place additional longevity and safety focused test requirements on wheel and brake suppliers. Added requirements, such as fail safe design verification(s), testing to failure of uncorroded and corroded wheels, and mandated use of overpressurization protection devices (in addition to fuses) have been incorporated. With the introduction of longer life tires, the demands on wheels and components, such as wheel bearings, have further intensified as landings between inspections at tire overhauls have increased. In addition, airplane manufacturers should be involved in wheel design, test, and manufacturer approvals to ensure that airplane specific needs have been addressed. For example, at least one airplane manufacturer specifies a missing wheel tie bolt requirement so that minimum equipment list (MEL) dispatch relief can be provided for a limited number of cycles. As a second example, individual airplane manufacturers may specify intensified wheel load and/or test requirements to account for various tire failure modes on multi-wheeled landing gear truck and other airplane landing gear configurations. Therefore, continued airworthiness of a particular wheel/tire or wheel/brake/tire assembly is usually demonstrated by compliance with airplane manufacturer requirements and not TSO minimum standards.
(d) Separate wheel applicable TSO and ( if applicable) airplane manufacturer tests are performed with radial and bias tires due to different tire-to-wheel interface loading and resultant wheel stress pattern and deflection/clearance differences. For braked wheels, separate wheel and brake assembly tests are performed with radial and bias tires due to differences in tire energy absorption that may be encountered. Some significant differences in wheel life have been reported by one wheel and brake supplier in one case using bias tires from different tire manufacturers. Typically, however, wheel/tire, or wheel/brake/tire assembly tests using different manufacturer bias tires have not been shown to be necessary.
(e) Wheel bearings in transport category airplane wheels should be qualified as part of the wheel assembly. Industry experience indicates that qualification of a specific manufacturer’s bearing(s) in a given wheel assembly is required to assure proper performance and airworthiness. Standard part bearing assemblies are not acceptable unless performance can be demonstrated during wheel qualification testing. It has been reported that roller end scoring is the most common mechanism leading to bearing failures on the airplane.
(f) Improved wheel bearing grease and bearing preload retention means have also been introduced on some recent airplanes to increase wheel bearing longevity in the severe
157 landing gear system environments and account for longer tire lives/ increased intervals between tire changes and inspections. Bearing wheel grease recommendations are usually specified in wheel and brake supplier CMMs. Intermixing of bearing cups/rollers/cones from different bearing manufacturers is not recommended on some large transport category airplane wheels due to different roller end scoring resistance capabilities and other often subtle differences.
c. Procedures/Method of Compliance . Due to the unique and critical nature of wheel, and wheel and brake designs, and historical airplane and personnel safety problems that have been experienced, compliance should only be approved upon successful completion of applicable TSO tests and any additional airplane manufacturer tests. Guidance should also be solicited from the original wheel and brake supplier(s)/TSO holders on any replacement part(s) to assure that continued airworthiness is not degraded.
54. Tires - § 25.733 [Reserved]
55. Brakes - § 25.735
a. References .
(1) TSO-C135a “Transport Airplane Wheels and Wheel and Brake Assemblies,” dated July 1, 2009.
(2) AC 21-29C, Change 1, “Detecting and Reporting Suspected Unapproved Parts,” dated July 22, 2008.
b. Explanation .
(1) Background .
(a) The original objective of § 25.735 (formerly § 4b.337) developed from a study to define a reasonable brake life for operational landings. This element is still retained in current § 25.735(f), which requires substantiation that the brakes have the ability to absorb the energy resulting from “operational landings at maximum landing weight,” and is directly related to the brake energy qualification tests of TSO-C135a . The two methods for compliance, stated in § 25.735(f), are also used in TSO-C135a for the determination of KEDL, the design landing kinetic energy rating for 100 stops at a deceleration rate of not less than 10 ft/sec2. The initial wear state of the brakes for the testing used to determine KEDL may be selected by the applicant.
Any condition representative of service use, including new, and that satisfies TSO-C135a (or acceptable equivalent), may be used. Section 25.735(f) makes reference to “operational landings” and “in the landing configuration,” which, in the context of modern transport airplane designs, could encompass a number of trailing edge flap positions and related landing speeds.
The actual “landing configuration” for showing compliance with § 25.735(f) is left to the option of the applicant, but should be selected on the basis of its being the predominant landing configuration expected in operational service.
(b) Over the years, as experience was gained in establishing takeoff and landing field lengths, it became evident that the rejected takeoff was critical in determining overall brake capability and could dictate aircraft maximum gross weight for dispatch.
(c) Investigation of an RTO overrun accident, in which 80 percent of the brakes on the subject airplane were at or very near their completely worn state, brought about the need to consider the effect of brake wear state on: 1) energy absorption capability, and 2) stopping capability. As a result, the FAA issued a series of specific airworthiness directives for the existing fleet of transport category airplanes to establish brake wear limits such that the brakes would be capable of absorbing a maximum energy absorption RTO in the fully worn state. The FAA also initiated rulemaking to address energy absorption capability and stopping distance with fully worn brake for future airplane types. The resulting rule, amendment 25-92, added:
1 A requirement in § 25.735(h) for the maximum rejected takeoff kinetic energy capacity rating of the aircraft brakes to be determined with the brakes at 100 percent of the allowable wear limit;
2 A new requirement in § 25.109(i) for the maximum kinetic energy rejected takeoff flight test demonstration to be conducted using brakes that have not more than 10 percent of their allowable wear range remaining; and
3 A general performance requirement under § 25.101(i) that requires the accelerate-stop and landing distances of §§ 25.109 and 25.125, respectively, to be determined with all wheel brake assemblies at the fully worn limit of their allowable wear range.
(2) Approval (§ 25.735(a)) . In accordance with § 21.305, a wheel and brake assembly may be approved under a TSO issued in accordance with subpart O to 14 CFR part 21; in conjunction with the type certification procedures for the airplane; or in any other manner approved by the Administrator, all of which are normally subject to flight test evaluation. (See TSO-C135a for additional guidance.)
(3) Brake tests (§ 25.735(b)) . If the applicant desires to make the maximum possible use of the brakes in establishing the landing distance, and if the contribution of the brakes to the total deceleration is relatively large, the brake system should be designed to permit the application of slightly less than half the braking deceleration developed under the conditions specified in this paragraph. The following dual system is recommended: dual wheel elements (drums or disc units), transmitting elements, power sources, master cylinders, etc., connected to a single pedal on each rudder pedal, such that the failure of any single one of these would leave half the total braking capacity symmetrically disposed about the plane of symmetry of the airplane. With such a system, it should be possible to show compliance with § 25.735(b) by means of calculations based upon the test data necessary to establish the landing distance, plus the brake data calculated by the airplane manufacturer.
(4) If the system is designed so that , under conditions herein specified, appreciably less than half the total braking capacity remains, or if the remaining capacity is asymmetrically disposed, tests should be conducted to determine that half the mean deceleration may, in fact, be
159 developed, and/or that the airplane may be safely controlled directionally while doing so. In order to obtain a minimum landing distance under § 25.125 and at the same time meet the deceleration requirements of § 25.735(b) in the event of failure of the normal brake system, it is common practice to provide an alternate brake system. When hydraulic (or pneumatic) brakes are used in the normal brake system, this alternate means usually consists of a duplicate hydraulic or pneumatic brake system and is commonly referred to as the “emergency brake system.” The following items should be considered in the design of such systems:
(a) Relationship between normal and emergency brake systems . The systems for actuating the normal brake and the emergency brake should be separated so that a failure (i.e., hydraulic leakage of fluid) from one system will not render the other system inoperative. A hydraulic brake assembly may be common to both the normal and emergency brake systems if it is shown that the leakage of hydraulic fluid resulting from failure of the sealing elements in the brake assembly would not reduce the airplane’s braking effectiveness below that specified in § 25.735(b).
(b) Brake controls (§ 25.735(c)) . General brake control force and operation should be noted throughout the flight test program to determine that they are satisfactory.
(c) Brake control valves . In the normal brake systems of all airplanes, the brake valves should be of a type such that the pilots may exercise variable control of the pressure to the brakes. The foregoing provision need not necessarily apply to the emergency brake system, although obviously such a provision would be desirable. Flight tests should be conducted to determine that the normal and emergency brake systems fulfill the requirements of § 25.231.
(d) Parking brake control (§ 25.735(d)) . A demonstration should be made to determine that sufficient braking is provided with the parking brake to prevent the airplane from rolling on a paved, dry, level runway (or any suitable level hard surface) while maximum takeoff power is applied on the most critical engine, with the airplane loaded to maximum ramp weight at aft center of gravity (or the airplane loaded to a weight-c.g. combination that prevents the wheels from sliding). In the case of propeller-driven airplanes, the effects of propeller wash and engine/propeller torque should also be considered in determining the critical engine. Because the resultant thrust vector can be at an angle to a propeller’s axis of rotation, one engine/propeller may be more critical than its counterpart on the opposite wing, particularly if all propellers turn in the same direction.
(e) Antiskid devices (§ 25.735(e)) .
1 In addition to meeting acceptable industry specifications, the installation of the antiskid device should comply with the requirements specified in paragraphs (2) and (3) below. The antiskid device and its installation will be approved for use on transport airplanes when the tests specified in the procedures section (see paragraph 55c of this AC) have been satisfactorily demonstrated.
2 Data required.
160 (aa) An engineering evaluation of the antiskid installation as installed on the airplane, including all necessary components, should be conducted. This analysis and complete descriptive data should be submitted to the FAA. The data should include hydraulic and electrical schematic diagrams of the installation, assembly drawings of antiskid system units, functional hazard analyses, software and hardware substantiations for electronic components, test results or stress analysis substantiating structural strength of attachments and modification of the axle or other structural members, installation drawings, recommended instructions pertaining to installation, maintenance and operation, and analysis of flight test data and results. Schematic drawings should refer to all units in the normal and emergency brake systems. The engineering evaluation should also assure that the antiskid system does not cause undesirable yaw characteristics.
(bb) The engineering evaluation should account for a bounce condition wherein the wheels may leave the runway after the brakes have been applied, for a condition wherein the wheels stay on the runway but the oleos are extended (if the system utilizes landing gear oleo compression in its operation), and for a condition in which the wheels of one main gear may not be in contact with the runway for a considerable time while the wheels of the other main gear are firmly on the runway. If the antiskid installation incorporates the “landing with brake pedals depressed” feature, then this type of operation should also be considered.
(cc) It should be shown that the brake cycling frequency imposed by the antiskid installation will not result in excessive loads on the landing gear.
(dd) The antiskid equipment should ensure satisfactory braking action on wet and snow/ice-covered runways, as well as on dry, hard-surfaced runways, without additional antiskid adjustments.
3 Systems requirements . The entire brake system (including both the basic brake system and the antiskid system) should conform to § 25.735. The single failure criterion of § 25.735 should be extended to include the antiskid system. The following should be considered for approval of antiskid systems:
(aa) In the event of a probable malfunction within the antiskid system that would result in loss of the antiskid feature on one or more brake units, those brake units affected should automatically revert to normal braking.
(bb) A means should be provided so that the pilot or copilot can readily deactivate the antiskid system to prevent loss of all braking capability due to antiskid system malfunctions.
(cc) A failure mode and effects analysis (FMEA) should be provided to evaluate the effect of failures of the antiskid and associated systems (tires, wheels, brakes, etc.)
on airplane safety and braking performance. For antiskid systems utilizing “locked wheel protection” circuits, this analysis should include a determination that assures mechanical faults, which could permit the affected axle to continue rotation (e.g., blown tires, broken wheel rims), will not significantly reduce the braking effectiveness of any cross-coupled wheel positions with
161 intact wheels and tires. A tire or wheel failure that results in the loss of additional tires or wheels, either from shrapnel impingement or ensuing overload, should be considered a single failure event.
(f) Wheel Fuse Plugs .
1 The hazardous condition of exploding tires and wheels associated with high energy emergency stopping conditions has been greatly alleviated by the installation of wheel fuse plugs. These plugs relieve the tire pressure when the wheel temperatures approach a dangerous limit. The effectiveness of these plugs in preventing hazardous tire blowouts should be demonstrated during an RTO test where the brake energy to be absorbed exceeds the maximum landing energy, but not the RTO energy, and the fuse plug releases, thereby deflating the tire(s) before blowout.
2 An improperly designed blowout plug that allows premature or unwanted release of tire pressure during takeoff or landing could also constitute a hazardous condition.
Such a situation would most probably arise during a takeoff from a quick turn-around type of airline operation. Fuse plug integrity should be demonstrated by conducting a maximum landing brake energy (which then becomes the quick turn-around chart limit for the AFM) test, which should not result in a fuse plug release.
3 Most turbojet transport airplanes have been able to demonstrate wheel blowout plug integrity at a maximum energy level in accordance with the procedures outlined in this section. More restrictive operational limitations (e.g., runway slope and tailwind values) have been imposed to stay within this maximum energy level demonstrated for wheel blowout plug integrity. With the advent of requests to increase the maximum landing weights, and to eliminate these restrictive operational limits, it has been considered acceptable to remove the pertinent restrictions and operational limitations and substitute in their place a chase-around chart as a limitation in the AFM. This chart will permit determining whether or not a critical energy level has been exceeded for the operating conditions of altitude, temperature, runway slope, tailwind, and landing weight. When the critical value is exceeded, a statement should be placed in the limitations section of the AFM to require that following a landing under such operating conditions, the airplane must remain on the ground a certain length of time prior to taxiing out for takeoff. This length of time will be the time to reach peak wheel temperatures (appropriate to the blowout plug location), plus 15 minutes. In lieu of the AFM fuse plug limitation chart, an alternate method for determining limit operational landing energy, such as a brake temperature limit, can be considered for approval.
4 The wear level of the brakes used for the chase-around chart discussed in the preceding paragraph should be selected by the applicant. Service experience indicates that the conservatism contained in the method of determining the turnaround time limits is adequate to allow these limits to be determined using new brakes.
5 In the case where it is possible to demonstrate wheel blowout plug integrity at a maximum energy level in accordance with the procedures in this section and without imposing certain restrictions on the operational limitations (e.g., runway slope and