AC 25-7D CHG 1
produce a 30-lb control force, or reach the trim limit, whichever occurs first.
10.3.2.1.2 Section 25.255(a)(2).
The maximum mistrim that can be sustained by the autopilot while maintaining level flight in the high-speed cruising condition. The highspeed cruising condition corresponds to the speed resulting from maximum continuous power or thrust, or VMO/MMO, whichever occurs first. Maximum autopilot mistrim may be a function of several variables, and the degree of mistrim should therefore correspond to the conditions of test. 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. If the autopilot is equipped with an auto-trim function, then the amount of mistrim that can be sustained, if any, will generally be small. If there is no auto-trim function, consideration should be given to the maximum amount of out-of-trim that can be sustained by the elevator servo without causing autopilot disconnect.
10.3.2.2 Section 25.255(b) establishes the basic requirement to show positive maneuvering stability throughout a specified acceleration envelope at all speeds to VFC/MFC, and the absence of longitudinal control force reversals throughout that acceleration envelope at speeds between VFC/MFC and VDF/MDF. (Later subsections (d) and (e) recognize that buffet boundary and control force limits will limit the acceleration actually reached; this does not account for Mach trim gain, etc.)
10.3.2.2.1 The out-of-trim condition for which compliance must be shown with § 25.255(b) is specified in § 25.255(a). For the initial trimmed condition before applying the mistrim criteria, the airplane should be trimmed at:
For speeds up to VMO/MMO, the particular speed at which the demonstration is being made; and For speeds higher than VMO/MMO, VMO/MMO.
10.3.2.2.2 Section 25.255(b)(2) appears to indicate that unstable airplane characteristics would be satisfactory, regardless of the character of the primary longitudinal control force as load factor is increased, as long as the force did not reverse (e.g., from a pull to a push). While such criteria may have merit for evaluating airplanes when starting the maneuver from a trimmed condition, it can be shown that this provides a poor specification for evaluating an airplane’s maneuvering characteristics when starting the test from the specified mistrimmed condition. For example, an airplane would be deemed to have unacceptable characteristics with a nose-up mistrim, if while relaxing the large initial elevator push force to increase the load factor to the specified value, the elevator force just happened to cross through zero to a slight pull force at 10-12
AC 25-7D CHG 1 one load factor, and then back through zero to a push force at a higher load factor. Such an airplane’s characteristics are clearly superior to one that has a severe elevator force slope reversal, during the same maneuver, but never reaches a zero elevator force condition as the load factor is increased. A literal interpretation of § 25.255(b)(2) would find this airplane to be compliant, while finding the preceding airplane non-compliant because it had a slight reversal of the primary longitudinal control force.
10.3.2.2.3 Section 25.255(b)(2) should be interpreted to mean that the primary longitudinal control force, for load factors greater than 1.0, may not be less than that used to obtain the initial 1 g flight condition. This is illustrated in figure 10-2 below. Slight control force reversals, as discussed in paragraph 10.3.2.2.1 above will be permitted for speeds between VFC/MFC and VDF/MDF only if:
No severe longitudinal control force slope reversals exist;
Any pitching tendency (uncommanded changes in load factor) should be mild and readily controllable; and The airplane’s pitch response to primary longitudinal control should be predictable to the pilot.
Figure 10-2. Mistrimmed Maneuvering Characteristics: Speeds Between VFC/MFC and VDF/MDF 10-13
AC 25-7D CHG 1
10.3.2.3 Section 25.255(c) requires that the investigation of maneuvering stability (§ 25.255(b)) include all attainable acceleration values between –l g and +2.5 g. Sections 25.333(b) and 25.337, to which it refers, limit the negative g maximum to 0 g at VD. Section 25.251 further limits the g to that occurring in probable inadvertent excursions beyond the buffet onset boundary at those altitudes where buffet is a factor.
10.3.2.4 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 +2 g agrees with predicted data, extrapolation to -1 g and 2.5 g should be allowed.
10.3.2.5 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 10.3.2.4 above.
10.3.2.6 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 -1 g to +1 g range. It limits the entry speed to avoid exceeding VDF/MDF.
10.3.2.7 Section 25.255(f) requires that, in the out-of-trim condition specified in § 25.255(a), it be possible to produce at least 1.5 g 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 (CG, VDF/MDF, altitude, etc.) of the airplane should be restricted to a value where 1.5 g is attainable. If trim must be used for the purpose of obtaining 1.5 g, it must be shown to operate with the primary control surface loaded to the least of three specified values.
10.3.2.7.1 The force resulting from application of the pilot limit loads of § 25.397 (300 lbs).
10.3.2.7.2 The control force required to produce 1.5 g (between 125 and 300 lbs).
10.3.2.7.3 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.
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AC 25-7D CHG 1
10.3.3 Procedures.
10.3.3.1 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.
10.3.3.2 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 dynamic pressure (q) and Mach limits.
10.3.3.3 The ability to move the primary controls (including trim), when loaded, should be considered prior to the tests.
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CHAPTER 11. STRUCTURE [RESERVED] 11-1
CHAPTER 12. DESIGN AND CONSTRUCTION: GENERAL [RESERVED] 12-1
CHAPTER 13. DESIGN AND CONSTRUCTION: CONTROL SURFACES [RESERVED] 13-1
CHAPTER 14. DESIGN AND CONSTRUCTION: CONTROL SYSTEMS
14.1 General—§ 25.671.
14.1.1 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.
14.1.2 Procedures.
14.1.2.1 In accordance with § 25.671(d), the airplane must be controllable when all engines fail. Compliance should be shown for each approved configuration. The airplane should remain controllable following the failure of the engines. Reconfiguration, if possible with all engines failed, is permitted. Any such reconfiguration should be included in the AFM operating procedures for an all engines failed condition.
14.1.2.2 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.
14.1.2.3 Past approaches to showing compliance with § 25.671(d) have been to show that the airplane 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.
14.1.2.4 For airplanes with fully powered or electronic flight control systems, the non-normal procedures section of the approved AFM should contain 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.”
14.2 Flap and Slat Interconnections—§ 25.701.
14.2.1 Explanation.
In accordance with § 25.701(a), if the wing flaps are not mechanically interconnected, tests or analysis should be conducted to show that the airplane has safe flight characteristics with asymmetric flap or slat deployment.
14-1
14.2.2 Reference.
See AC 25-22 for additional guidance.
14.2.3 Procedures.
Simulate appropriate flap and slat malfunctioning during takeoffs, approaches, and landings to demonstrate that the airplane is safe under these conditions. To be considered safe, adequate stall margins and controllability should be retained without requiring exceptional piloting skill or strength. Additionally, there should be no hazardous change in altitude or attitude during transition to the asymmetric condition considering likely transition rates.
14.3 Takeoff Warning System—§ 25.703. [Reserved] 14-2
CHAPTER 15. DESIGN AND CONSTRUCTION: LANDING GEAR
15.1 Retracting Mechanism—§ 25.729.
15.1.1 Explanation.
None.
15.1.2 Procedures.
15.1.2.1 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:
15.1.2.1.1 VLO (the placard airspeed) in the cruise configuration at near 1 g flight and normal yaw angles; and
15.1.2.1.2 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 associated with engine-out flight and counteracting crosswinds of up to 20 knots.
15.1.2.2 The alternate extend system should be demonstrated at airspeeds up to VLO at near 1 g 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 the note under paragraph 15.1.2.1.2 above.)
15.1.2.3 Operation Test—§ 25.729(d).
15.1.2.3.1 The engine-out gear retraction time should be determined from flight tests with one engine at idle power or thrust and the operating engine(s) adjusted to provide the lowest thrust-to-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.
15.1.2.3.2 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.
15-1
15.1.2.4 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.
15.2 Wheels—§ 25.731.
15.2.1 References.
TSO-C135a, Transport Airplane Wheels and Wheel and Brake Assemblies, dated July 1, 2009.
Paragraph 15.4.2.5 of this AC, Wheel Fuse Plugs.
Paragraph 15.4.10 of this AC, Wheel Fuse Plug Integrity.
AC 21-29D.
15.2.2 Explanation.
15.2.2.1 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. Paragraphs 15.4.2.5 and 15.4.10 of this AC provide insight into the criticality of proper fuse function to support airworthiness. As demands 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 rollon-rim requirement to increase robustness in wheel flange areas.
15.2.2.2 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 supplier’s 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.
15.2.2.3 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 15-2
verification(s), testing to failure of uncorroded and corroded wheels, and mandated use of over pressurization 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.
15.2.2.4 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.
15.2.2.5 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.
15.2.2.6 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 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 15-3