(i) The evaluation should proceed to a g level that will allow recovery to be accomplished near +2.5g, unless sufficient buffet or other phenomena (natural, artificial, or a combination) of such intensity exists that is a strong and effective deterrent to further pilot application of nose-up longitudinal control force (as in § 25.201(d)(2)) so that there is no danger of exceeding the airplane limit load factor (Ref. § 25.143(b)).
NOTE: A strong and effective deterrent is analogous to that required for stall identification; stick shaker or stall warning buffet are not considered to be an adequate end point for these tests.
(ii) Any pitching tendency ( uncommanded changes in load factor) shall be mild and readily controllable.
(iii) Sufficient control s hall be available to the pilot, through unreversed use of only the primary longitudinal control, to effect a prompt recovery to +1g flight from the load factors described herein.
(iv) The airplane's pitch response to primary longitudinal control sha ll be predictable to the pilot.
(B) Experience has shown that maneuvering evaluations conducted at the highest Mach and the highest weight and altitude (W/delta) combination may not necessarily produce the most critical results. Equally important is the character of the buffet buildup (e.g., slowly increasing or rapid rise, and the g at which it starts).
Conditions associated with buffet onset near 2g at Mach numbers below M MO have sometimes yielded the most critical characteristics. Therefore, a sufficient spread of conditions should be evaluated.
FIGURE 31-1. MANEUVERING CHARACTERISTICS Speeds Up to the Lesser of V FC or MFC or Buffet Onset at 1G * These characteristics are satisfactory only in accordance with paragr aphs b(5)( i)(A)(1) and (2).
FIGURE 31-2. MANEUVERING CHARACTERISTICS Speeds Between Figure 31-1 and V DF or MDF * These characteristics are satisfactory only in accordance with paragraphs b(5)( i)(A)(1) and (2).
32. HIGH SPEED CHARACTERISTICS - § 25.253 .
a. Explanation.
(1) The maximum flight demonstrated dive speed, V DF/MDF, is used when establishing V MO/MMO and the associated speed margins under the provisions of § 25.1505. Both V MO and MMO are then evaluated during flight tests for showing compliance with § 25.253.
(2) The pitch upset described in § 25.335(b), as amended by Amendment 25-23, or § 25.1505, prior to Amendment 25-23, is for design criteria only. The operational upsets expected to occur in service for pitch, roll, yaw, and combined axis upsets are covered under § 25.253.
(3) In general, the same maneuvers should be accomplished in both the dynamic pressure and Mach critical ranges. All maneuvers in either range should be accomplished at thrust and trim points appropriate for the specific range. It must be realized that some maneuvers in the Mach range may be more critical for some airplanes due to drag rise characteristics, and that at high altitudes a lower gross weight may be required to achieve the maximum approved operating altitude and Mach/airspeed conditions.
(4) The airplane's handling characteristics in the high speed range should be investigated in terms of anticipated action on the part of the flightcrew during normal and emergency conditions.
(5) At least the following factors should be considered in determining the necessary flight tests:
(i) Effectiveness of longitudinal control at V MO/MMO and up to VDF/MDF.
(ii) Effect of any reasonably probable mistrim on upset and recovery.
(iii) Dynamic and static stability.
(iv) The speed increase that results from likely passenger movement when trimmed at any cruise speed to VMO/MMO.
(v) Trim changes resulting from compressibility effects.
(vi) Characteristics exhibited during recovery from inadvertent speed increase.
(vii) Upsets due to vertical and horizontal gusts (turbulence).
(viii) Speed increases due to horizontal gusts and temperature inversions.
(ix) Effective and unmistakabl e aural speed warning at V MO plus 6 knots, or MMO plus 0.0lM.
(x) Speed and flightpath control during application of deceleration devices.
(xi) The manageability of control forces resulting from the application of deceleration devices.
(6) Section 25.1505 states that the speed margin between V MO/MMO, and VD/MD or VDF/MDF, as applicable, "may not be less than that determined under § 25.335(b) or found necessary during the flight tests conducted under § 25.253." Historically, this speed margi n allowance for atmospheric variations (e.g., horizontal gusts, penetration of jet streams and cold fronts), instrument errors, and airframe production tolerances has been interpreted as .05M in the Mach limited portion of the flight envelope. However, studies conducted by industry during the 1980s have shown that for a conventional aircraft, atmospheric disturbances resulted in larger airspeed variations than accounted for by the .05M minimum margin. Consequently, this speed margin was increased to .07 Mach by Amendment 25-91, with an allowance for reducing the margin to the old value of .05M if justified by a rational analysis. In establishing the minimum speed margins during type certification programs, the factors outlined in paragraph (5), above, should also be considered in addition to the items listed below:
(i) Increment allowance for production tolerances in airspeed systems (0.005M), unless larger differences are found to exist.
(ii) Increment allowance for production tolerances of overs peed warning error (0.0lM), unless larger tolerances or errors are found to exist.
(iii) Increment allowance ∆M due to speed overshoot from M MO, established by upset during flight tests in accordance with § 25.253, should be added to the values for production differences and equipment tolerances, and the minimum acceptable combined value should not be less than .07M between M MO and MD/MDF. The value of MMO then should not be greater than the lowest value obtained from each of the following equations and from § 25.l505:
MMO ≤ MD/MDF - ∆M - .005M - .0lM or MMO ≤ MD/MDF - .07M NOTE: The combined minimum increment may be reduced from .07M to as small as .05M if justified by the rational analysis used to show compliance with § 25.335(b)(2).
(iv) At altitudes where V MO is limiting, the increment allowance for production differences of airspeed systems and production tolerances of overspeed warning errors are 3 and 6 knots, respectively, unless larger differences or errors are found to exist.
(v) Increment allowance ∆V due to speed overshoot from V MO, established by upset during flight tests in accordance with § 25.253, should be added to the values for production differences and equipment
tolerances. The value of V MO should not be greater than the lowest obtained from the following equation, and from § 25.l505:
VMO ≤ VD/VDF - ∆V - 3 knots - 6 knots (production differences) (equipment tolerances)
(vi) For an airplane with digital interface between the airspeed system and the ove rspeed warning system, the production tolerance for the warning system may be deleted when adequately substantiated.
b. Regulations Affected . These criteria refer to certain provisions of Part 25 of the Federal Aviation Regulations (FAR). They may also be used in showing compliance with the corresponding provisions of the former Civil Air Regulations (CAR) in the case of airplanes to which those regulations are applicable. Other affected FAR are as follows:
Section 25.175(b) Demonstration of stati c longitudinal stability.
Section 25.251 Vibration and buffeting.
Section 25.253 High-speed characteristics.
Section 25.335(b) Design dive speed, V D.
Sections 25.1303(b)(1) and (c) Flight and navigation instruments.
Section 25.1505 Maximum operating limit speed.
c. Procedures . Using the speeds VMO/MMO and VDF/MDF determined in accordance with §§ 25.1505 and
25.251, respectively, and the associated speed margins, the airplane should be shown to comply with the high-speed characteristics of § 25.253. Unless otherwise stated, the airplane characteristics should be investigated at the most critical speed up to and including V MO/MMO, and the recovery procedures used should be those selected by the applicant, except that the normal acceleration during recovery should be no more than l.5g (total). Testing should be conducted with the c.g. at the critical position and generally perpendicular to local wind aloft.
(1) Center of Gravity Shift . The airplane should be upset by the center of gravity shift corresponding to the forward movement of a representative number of passengers (and/or serving carts) depending upon the airplane interior configuration. The airplane should be permitted to accelerate until 3 seconds after effective overspeed warning.
(2) Inadvertent Speed Increase . Simulate an evasive control application when trimmed at V MO/MMO, by applying sufficient forward force to the pitch control to produce 0.5g (total) for a period of 5 seconds, after which recovery should be effected at not more than l.5g (total).
(3) Gust Upset . In the following three upset tests, the values of displacement should be appropriate to the airplane type and should depend upon airplane stability and inertia characteristics. The lower and upper limits should be used for airplanes with low and high maneuverability, respectively.
(i) With the airplane trimmed in wings-level flight, simulate a transient gust by rapidly rolling to the maximum bank angle appropriate for the airplane, but not less than 45 degrees nor more than 60 degrees. The rudder and longitudinal control should be held fixed during the time that the required bank is being attained. The rolling velocity should be arrested at this bank angle. Following this, the controls should be abandoned for a minimum of 3 seconds after overspeed warning or 10 seconds, whichever occurs first.
(ii) Perform a longitudinal upset from normal cruise. Airplane trim is determined at V MO/MMO using power/thrust required for level flight but with not more than maximum continuous power/thrust. This is followed by a decrease in speed after which an attitude of 6 -l2 degrees nose down, as appropriate for the airplane type, is attained with the power/thrust and trim initially required for V MO/MMO in level flight. The airplane is
permitted to accelerate until 3 seconds after overspeed warning. The force limits of § 25.143(c) for short term application apply.
(iii) Perform a two-axis upset, consisting of combined longitudinal and lateral upset s. Perform the longitudinal upset, as in paragraph (ii) above, and when the pitch attitude is set, but before reaching V MO/MMO, roll the airplane l5-25 degrees. The established attitude should be maintained until 3 seconds after overspeed warning.
(4) Leveling Off from Climb . Perform transition from climb to level flight without reducing power below the maximum value permitted for climb until 3 seconds after overspeed warning. Recovery should be accomplished by applying not more than 1.5g (total).
(5) Descent from Mach Airspeed Limit Altitude . A descent should be performed at the airspeed schedule defined by M MO and continued until 3 seconds after overspeed warning occurs, at which time recovery should be accomplished without exceeding 1.5g (total).
(6) Application of Deceleration Devices . It should be demonstrated that extension of the airbrakes at high speeds will not result in excessive flight path deviation. With the airplane trimmed for level flight at VMO/MMO, extension of the airbrakes to any position that may be selected by the pilot's control, at speeds above VMO/MMO but not so high that V DF/MDF would be exceeded during the maneuver, should not result in:
(i) A positive normal load factor greater than 2.0 with the stick free; or
(ii) A nose down pitching moment that requires more than a 20 pound longitudinal control force to maintain 1g flight.
33. OUT-OF-TRIM CHARACTERISTICS - § 25.255 .
a. Explanation. Certain early, trimmable stabilizer equipped jet transports experienced "jet upsets" that resulted in high speed dives. When the airplane was mistrimmed in the nose-down direction and allowed to accelerate to a high airspeed, it was found that there was insufficient elevator power to recover. Also, the stabilizer could not be trimmed in the nose-up direction, because the stabilizer motor stalled due to excessive airloads imposed on the horizontal stabilizer. As a result, a special condition was developed and applied to most Part 25 airplanes with trimmable stabilizers. With certain substantive changes, it was adopted as § 25.255, effective with Amendment
25-2. While these earlier problems seem to be generally associated with airplanes having trimmable stabilizers, it is clear from the preamble discussions to Amendment 25-42 that § 25.255 applies "regardless of the type of trim system used in the airplane." Section 25.255 is structured to give protection against the following unsatisfactory characteristics during mistrimmed flight in the higher speed regimes:
(1) Changes in maneuvering stability leading to overcontrolling in pitch.
(2) Inability to achieve at least l.5g for recovery from upset due to excessive control forces.
(3) Inability of the flightcrew to apply the control forces necessary to achieve recovery.
(4) Inability of the pitch trim system to provide necessary control force relief when high control force inputs are present.
b. Reference Regulation . Section 25.255
c. Discussion of The Regulation .
(1) Section 25.255(a) is the general statement of purpose. Maneuvering stability may be shown by a plot of applied control force versus normal acceleration at the airplane center of gravity. Mistrim must be set to the greater of the following:
(i) Section 25.255(a)(l). A 3-second move ment of the longitudinal trim system at its normal rate for the particular flight condition with no aerodynamic load. Since many modern trim systems are variable rate
systems, this subsection requires that the maneuver condition be defined and that the no-load trim rate for that condition be used to set the degree of mistrim required. For airplanes that do not have power-operated trim systems, experience has shown a suitable amount of longitudinal mistrim to be applied is that necessary to produce a 30 pound control force, or reach the trim limit, whichever occurs first.
(ii) 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 high speed cruising condition corresponds to the speed resulting from maximum continuous power or thrust, or V MO/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.
(2) Section 25.255(b) establishes the basic requirement to show positive maneuvering stability throughout a specified acceleration envelope at all speeds to V FC/MFC, and the absence of longitudinal control force reversals throughout that acceleration envelope at speeds between V FC/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.)
(i) Section 25.255(b)(2), taken at face value, indicates 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 an anomalous specification for evaluation of an airplane's maneuvering characteristics when starting the test from the specified mistrimmed condition. For example, an airplane would 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 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 pass this airplane while failing the preceding airplane because it had a slight reversal of the primary longitudinal control force.
(ii) 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 1g flight condition. This is illustrated in Figure 33-1 and is consistent with the criteria depicted in Figure 31-2 for in-trim maneuver stability.
Slight control force reversals, as discussed in paragraph ( i), above, will be permitted for speeds between V FC/MFC and VDF/MDF only if:
(A) No severe longitudinal control force slope reversals exist;
(B) Any pitching tendency ( uncommanded changes in load factor) shall be mild and readily controllable; and (C) The airplane's pitch response to prima ry longitudinal control shall be predictable to the pilot.
FIGURE 33-1. MISTRIMMED MANEUVERING CHARACTERISTICS Speeds Between V FC/MFC and VDF/MDF
(3) Section 25.255(c) requires that the investigation of maneuvering stability (§ 25.255(b)) includ e 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 V D. 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 V DF/MDF.
(7) Section 25.255(f) requires that in the out-of-trim condition of paragraph (a) it must be possible to produce at least 1.5g during recovery from the overspeed condition of V DF/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.
(i) The force resulting from application of the pilot limit loads of § 25.397 ( 300 lbs.).
(ii) The control force required to produce 1.5g (between 125 and 300 lbs.).
(iii) 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 F S/g (normally a plot). Any standard flight test procedure that yields an accurate evaluation of F S/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 Airplane Flight Manual 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 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(4) of this AC and Advisory Circular 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, flig ht 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:
(i) VLO (the placard airspeed) in the cruise configuration at near 1g flight and normal yaw angles;
and
(ii) 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 V S1 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.
(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).
(i) The engine-out gear retraction time should be d etermined from flight tests with one engine at idle power 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.
(ii) 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) Technical Standard Order (TSO) C26c, "Aircraft Wheels and Wheel-B rake Assemblies, with Addendum I," dated May 18,1984, for additional guidance.
(2) Paragraph 55b(4)vi of this AC, Wheel Fuse Plugs.
(3) Paragraph 55c(7) of this AC, Wheel Fuse Plug Design.
(4) Advisory Circular 21-29A, " Detecting and Reporting Su spected Unapproved Parts," dated July 16, 1992.
b. Explanation.
(1) Background .
(i) 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 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 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.
(ii) 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.
(iii) 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.
(iv) 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.
(v) 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. In response to demands for more robust wheel bearing systems, due in part to longer life tires and increased numbers of landings between wheel and bearing inspections, at least one airplane manufacturer has required qualification testing of each wheel and specific wheel bearing supplier assembly to test requirements exceeding those in TSO-C26c and typical airplane manufacturer specifications. This has required suppliers of larger wheel and brake assemblies to develop special test capabilities for qualification of specific wheel bearing assembly systems (i.e.: bearing manufacturer, grease, seals and other means of grease retention). It has been reported that roller end scoring is the most common mechanism leading to bearing failures on the airplane.