(2) The net flight path shall have a positive slope at 1,500 feet above the airport where the landing is assumed to be made after failure of two engines.
(3) Fuel jettisoning shall be permitted if the Administrator finds that the operator has an adequate training program, proper instructions are given to the flight crew, and all other precautions are taken to ensure a safe procedure.
(4) The airplane’s weight at the point where the two engines are assumed to fail shall be considered to be not less than that which would include sufficient fuel to proceed to the airport and to arrive there at an altitude of at least 1,500 feet directly over the landing area and thereafter to fly for 15 minutes at cruise power and/or thrust.
(5) The consumption of fuel and oil after the engines are assumed to fail shall be that which is accounted for in the net flight path data shown in the Airplane Flight Manual.
40T.84 Landing limitations.
(a) Airport of destination. No airplane shall be taken off at a weight in excess of that which, in accordance with the landing distances shown in the Airplane Flight Manual for the elevation of the airport of intended destination and for the wind conditions anticipated there at the time of landing, would permit the airplane to be brought to rest at the airport of intended destination within 60 percent of the effective length of the runway from a point 50 feet directly above the intersection of the obstruction clearance plane and the runway. The weight of the airplane shall be assumed to be reduced by the weight of the fuel and oil expected to be consumed in flight to the airport of intended destination. Compliance shall be shown with the conditions of subparagraphs (1) and (2) of this paragraph. (See secs. 4T.123(b) amd 4T.743(b).)
(1) It shall be assumed that the airplane is landed on the most favorable runway and direction in still air.
(2) It shall be assumed that the airplane is landed on the most suitable runway considering the probable wind velocity and direction and taking due account of the ground handling characteristics of the airplane and of other conditions (i.e., landing aids. terrain, etc.). If full compliance with the provisions of this subparagraph is not shown, the airplane may be taken off if an alternate airport is designated which permits compliance with paragraph (b) of this section.
(b) Alternate airport. No airport shall be designated as an alternate airport in a dispatch release unless the airplane at the weight anticipated at the time of arrival at such airport can comply with the provisions of paragraph (a) of this section, provided that the airplane can be brought to rest within 70 percent of the effective length of the runway.
4. In lieu of section 43.11 of Part 43 of the Civil Air Regulations the following shall be applicable.
43T.11 Transport category airplane weight limitations. The performance data in the Airplane Flight Manual shall be applied in determining compliance with the following provisions:
(a) No airplane shall be taken off at a weight which exceeds the takeoff weight specified in the Airplane Flight Manual for the elevation of the airport and for the ambient temperature existing at the time of the takeoff. (See secs. 4T.123(a)(1) and 4T.743(a).)
(b) No airplane shall be taken off at a weight such that, allowing for normal consumption of fuel and oil in flight to the airport of destination and to the alternate airports, the weight on arrival will exceed the landing weight specified in the Airplane Flight Manual for the elevation of each of the airports involved and for the ambient temperatures anticipated at the time of landing. (See secs. 4T.123(a) (2) and 4T.743(a).)
(c) No airplane shall be taken off at a weight which exceeds the weight at which, in accordance with the minimum distances for takeoff scheduled in the Airplane Flight Manual,
compliance with subparagraphs (1) through (3) of this paragraph is shown. These distances shall correspond with the elevation of the airport, the runway to be used, the effective runway gradient, and the ambient temperature and wind component existing at the time of takeoff. (See secs. 4T.123(a)(3) and 4T.734(a).)
(1) The accelerate-stop distance shall not be greater than the length of the runway plus the length of the stopway if present.
(2) The takeoff distance shall not be greater than the length of the runway plus the length of the clearway if present, except that the length of the clearway shall not be greater than onehalf of the length of the runway.
(3) The takeoff run shall not be greater than the length of the runway.
(d) No airplane shall be operated outside the operational limits specified in the Airplane Flight Manual. (See secs. 4T.123(a) (4) and 4T.743(a).)
5. The following definitions shall apply:
(a) Clearway. A clearway is an area beyond the runway, not less than 500 feet wide, centrally located about the extended centerline of the runway, and under the control of the airport authorities. The clearway is expressed in terms of a clearway plane, extending from the end of the runway with an upward slope not exceeding 1.25 percent, above which no object nor any portion of the terrain protrudes, except that threshold lights may protrude above the plane if their height above the end of the runway is not greater than 26 inches and if they are located to each side of the runway.
NOTE: For the purpose of establishing takeoff distances and takeoff runs, in accordance with section 4T.117 of this regulation, the clearway plane is considered to be the takeoff surface.
(b) Stopway. A stopway is an area beyond the runway, not less in width than the width of the runway, centrally located about the extended centerline of the runway, and designated by the airport authorities for use in decelerating the airplane during an aborted takeoff: To be considered as such, a stopway must be capable of supporting the airplane during an aborted takeoff without inducing structural damage to the airplane. (See also sec. 4T.115(d) of this regulation.)
APPENDIX 2
HISTORICAL DEVELOPMENT OF ACCELERATE-STOP TIME DELAYS
1. Explanation.
a. As the rules pertaining to transport category airplane accelerate-stop distances have developed over the years, the interpretation of those rules, with regard to pilot actions, recognition times, and delay times, has also changed. The paragraphs below provide a historical perspective on the application of pilot recognition and delay times, with references to the applicable 14 CFR part 25 amendment number.
b. In accordance with the stipulations of §§ 21.101(a) and (b), the certification basis for a derivative airplane type may be the regulations specified in the type certificate of the basic airplane, depending upon the extent and nature of the changes to the airplane. For those cases where the original certification basis can be retained, the guidance presented in paragraphs 2a and b., below, may be used to show compliance with the appropriate regulations. (For more information related to determining a derivative airplane's certification basis, see FAA Order 8110.4A, "Type Certification Process.")
c. Regardless of the certification basis, the effects of brake wear state on energy absorption capability and stopping performance must be addressed. The responsible FAA Aircraft Certification Office (ACO) should be contacted to establish the applicable brake wear criteria for a specific airplane type.
2. Procedures .
a. Pre-Amendment 25-42 Accelerate-Stop Time Delays . Parts 1 and 25 of Title 14 of the CFR, prior to Amendments 1-29 and 25-42, respectively, defined V 1 as the critical engine failure speed. When this definition of V1 was applied to the accelerate-stop criteria of § 25.109 and the V 1 criteria of § 25.107(a)(2), the implication was that engine failure and engine failure recognition could occur simultaneously. It was recognized that this simultaneous occurrence could not be achieved in actual operations, and that defining V 1 as the engine failure speed resulted in a conflict with § 25.101(h), which requires allowance for time delays in execution of procedures. In order to resolve this conflict, V 1 was applied as the engine failure recognition speed, and appropriate time delays were developed for showing compliance with § 25.101(h). Figure 1 shows a pictorial representation of the accelerate-stop time delays that were considered acceptable for compliance with § 25.101(h) before Amendment 25- 42 (effective March 1, 1978). Paragraph b(1)(vii) of this appendix provides guidance on reflecting these time delays in the Airplane Flight Manual accelerate-stop distances.
FIGURE 1. ACCELERATE-STOP TIME DELAYS (Pre-Amendment 25-42)
(1) ∆trec = engine failure recognition time. The demonstrated time from engine failure to pilot activation of the first deceleration device, this action indicating recognition of the engine failure. For AFM data expansion purposes, in order to provide a recognition time increment that can be executed consistently in service, it has been found practical to use the demonstrated time or 1 second, whichever is greater. If the airplane incorporates an engine failure warning light, the recognition time includes the time increment necessary for the engine to spool down to the point of warning light activation, plus the time increment from light "on" to pilot action indicating recognition of engine failure.
(2) ∆ta1 = the demonstrated time interval between activation of the first and second deceleration devices.
(3) ∆ta2 = the demonstrated time interval between activation of the second and third deceleration devices.
(4) ∆t = a 1-second reaction time delay to account for in-service variations. If a command is required for another crewmember to actuate a deceleration device, a 2-second delay, in lieu of the 1-second delay, should be applied for each action. For automatic deceleration devices that are approved for performance credit for AFM data expansion, established times determined during certification testing may be used without the application of the additional time delays required by this paragraph.
(5) The sequence for activation of deceleration devices may be selected by the applicant in accordance with § 25.101(f). If, on occasion, the desired sequence is not achieved during testing, the test need not be repeated;
however, sufficient tests must be conducted to establish acceptable values of ∆ta.
(6) If additional devices are use d to decelerate the airplane, their respective demonstrated times, plus any additional required time delays, should be included until the airplane is in the full braking configuration.
(7) For the purpose of flight manual calculations, the 1-second del ay for each action may be added at the end of the total demonstrated time. Regardless of the manner in which the time delays are applied, the flight manual calculations must assume the airplane does not decelerate during the delay time increments.
b. Amendment 25-42 thru Amendment 25-91, Accelerate-Stop Time Delays . Amendment 25-42, effective March 1, 1978, introduced several new requirements affecting accelerate-stop distance determination. One of the most significant changes of Amendment 25-42 was the requirement to determine an all-engines-operating accelerate-stop distance to account for the many rejected takeoffs that were not the result of an engine failure.
Amendment 25-42 also introduced into the regulations an engine -failure speed, VEF; redefined the takeoff decision speed, V1; revised the accelerate -stop distance criteria to correspond to the V EF and V1 definitions; and added a 2- second time delay between V 1 and the first action to decelerate the airplane during which the airplane continues
accelerating with the operating engines at takeoff thrust. Figure 2 shows a pictorial representation of the acceleratestop time delays in accordance with the provisions of part 25, including Amendment 25-42:
FIGURE 2 ACCELERATE-STOP TIME DELAYS (Amendment 25-42 thru Amendment 25-91)
(1) ∆trec - § 25.107 defines the relationship between V EF and V1 as follows:
"VEF is the calibrated airspeed at which the critical engine is assumed to fail. V EF must be selected by the applicant, but may not be less than V MCG determined under §
25.149(e). V1 , in terms of calibrated airspeed, is the takeoff decision speed selected by the applicant; however, V 1 may not be less than V EF plus the speed gained with the critical engine inoperative during the time interval between the instant at which the critical engine is failed, and the instant at which the pilot recognizes and reacts to the engine failure, as indicated by the pilot's application of the first retarding means during accelerate-stop tests."
(2) Demonstrated engine failure recognition times less than 1 second should be carefully reviewed to assure the conditions under which they were obtained were representative of that which may reasonably be expected to occur in service. A sufficient number of demonstrations should be conducted using both applicant and FAA test pilots to assure that the time increment is representative and repeatable. The pilot's feet should be on the rudder pedals, not brakes, during demonstration tests.
(3) ∆ta1 = the demonstrated time interval between activation of the first and second deceleration devices.
(4) ∆ta2 = the demonstrated time interval between activation of the second and third deceleration devices.
(5) If a command is required for another crewmember to activate a deceleration device, a 1 -second delay, in addition to the delays specified in paragraphs (3) and (4) above, should be applied for each action. For automatic deceleration devices which are approved for performance credit for AFM data expansion, established system times determined during certification testing may be used. These established times cannot be assumed to
start until after the pilot action that triggers them; that is, they cannot be triggered before the first pilot action and, hence, cannot begin until at least 2 seconds after V 1.
(6) The sequence for activation of deceleration devices may be selected by the applicant in accordance with § 25.101(f). If, on occasion, the desired sequence is not achieved during testing, the test need not be repeated; however, sufficient tests must be conducted to establish acceptable values of ∆ta.
(7) Figure 2 shows a pictorial representation of how to apply time delays with up to three deceleration devices. If more than three devices are used to decelerate the airplane, the respective demonstrated time plus a 1-second reaction time delay should be included for each device beyond that represented pictorially in Figure 2 until the airplane is in the full braking configuration.
APPENDIX 3
CORRECTION OF AIR MINIMUM CONTROL SPEED TO STANDARD CONDITIONS The following analysis presents three methods of correcting a flight test derived value of air minimum control speed to standard conditions. These methods are applicable only to rudder deflection limited VMCA, for either jet or propeller driven airplanes. The effect of banking into the operating engine is accounted for, and the method will work with either fixed pitch or constant speed propellers, including the effects of windmilling drag. For rudder pedal force limited VMCA, see Appendix 4.
Theoretical Basis Given the static lateral/directional equations of motion for straight line, unaccelerated flight:
Fy∑ = 0 Cyb b⋅ + Cy aad d⋅ + Cy rrd d⋅ = CL · sin φ (1) Mx∑ = 0 Clb b⋅ + Cl aad d⋅ + Cl rrd d⋅ = 0 (2) Mz∑ = 0 Cnb b⋅ + Cn aad d⋅ + Cn rrd d⋅ = Cna (3) where:
CL = 295 2 ⋅ ⋅ W V Se Cna = 295 2 ⋅ ⋅ ⋅ ⋅ F l V S b na e e W = weight - lbs Ve = equivalent airspeed - kts S = wing area - ft2 Fna = asymmetric net thrust - lbs Fna = (Fn + Dw) for engine inoperative Fna = (Fn + Fi) for engine at idle Fn = net thrust of the operating engine - lbs Fi = idle engine net thrust - lbs Dw = windmill drag - lbs le = distance from aircraft center line to engine thrust line - ft b = wingspan - ft Constant Cn Method For the case where full rudder deflection is achieved, δr is a constant, and the system of equations can be resolved to an identity which shows that Cna is a linear function of CLsin φ.
Cna = A·CL ⋅ sinφ + B (4)
If it is assumed that test and standard day VMCA occur at the same angle of attack and bank angle, the asymmetric yawing moment coefficient will be constant, and V MCA can be corrected to standard conditions by the relationship:
VMCAs = VMCAt · F F na na s t for Turbojets VMCAs = VMCAt · THP THP s s t t ⋅ ⋅ s s 1 3 for Propeller driven where:
*THPs = Maximum AFM scheduled Brake/Shaft horsepower multiplied by sta ndard day propeller efficiency.
THPt = Test day brake/shaft horsepower where V MCA was achieved multiplied by test day propeller efficiency.
s s = Atmospheric density ratio at standard conditions.
s t = Atmospheric density ratio at test conditions.
*Windmilling SHP is not considered, because the current 14 CFR part 25 takeoff requirements for propeller driven airplanes result in such large performance penalties with a windmilling propeller that all part 25 turboprops to date have had autofeather installed.
Since both net thrust and shaft horsepower vary with speed, use of these equations will require an iterative solution.
Because this constant Cn method does not consider the effect on V MCA due to variations in bank angle, weight, sideslip angle, or adverse yaw, its use is limited to corrections of 5% or less in asymmetric net thrust or power.
For corrections beyond 5%, the relationship shown in equation (4) must be used, and enough flight test data must be obtained to define the correlation between Cna and CLsin φ.
Graphical Method In theory these data could be obtained by varying any combination of asymmetric thrust, airspeed, weight, and bank angle that would provide a representative variable set. However, since V MCA and stall speed are nearly coincident for most airplanes, there are some severe constraints on most of the variables. Typically, any reduction in maximum asymmetric thrust will cause V MCA to decrease below stall speed, and any increase in weight will cause stall speed to increase above VMCA; therefore, the only parameter that can reasonably be varied is bank angle.
To maximize the spread between stall and minimum control speed, V MCA tests are normally done at the lightest possible weight, at the maximum allowable asymmetric thrust (even with a short duration overboost, if the engine manufacturer will agree). At typical test altitudes (2000-3000 feet) and prototype gross weights, it will usually still not be possible to define V MCA with the full 5° bank, because of stall buffet.
To obtain the data necessary for extrapolation to the 5° bank limit, and to maximum asymmetric thrust, testing at three bank angles is required for the definition of the Cna vs CL sin φ relationship. These data should be obtained by shutting down the critical engine (normally the left), setting maximum allowable thrust on the operative engine, and slowing down while maintaining constant heading until full rudder deflection is achieved. The first point, a wings level condition, is easy to set up, and results in a speed well above stall buffet. A second point, at zero sideslip, will be achieved at approximately 2° - 3° bank (flown with a yaw string, or instrumented sideslip vane) and will provide an intermediate speed, still above buffet. The third data point is flown with as much bank angle that can be used without excessive buffeting (no more than would be accepted as the minimum level of stall warning). If necessary, an additional point can be obtained by banking 2° - 3° into the inoperative engine.
To use this method, instrumentation is necessary for the determination of net thrust/shaft horsepower, and an accurate calibrated airspeed system is required, as well as engine/propeller charts for windmill drag, charts for propeller efficiency, and the ability to measure bank angle to at least a tenth of a degree.
Data obtained using this method with a typical business jet and a large jet transport are shown in the following two graphs:
2-ENGINE BUSINESS JET 0
0.01
0.02
0.03
0.04
0.05
0.06 -0.04 0 0.04 0.08 0.12 0.16 CLsinφ Cna Kβ=0.20 4-ENGINE TRANSPORT 0
0.01
0.02
0.03
0.04
0.05
0.06
0.07
0.08 0 0.05 0.1 0.15 0.2 0.25 CLsinφ Cna Kβ=0.14
These plots represent the capability of the AIRFRAME to produce yawing moment by a combination of rudder deflection (full, in this case), and the sideslip which results from the bank angle. In order to determine the limiting condition for VMCA, it is necessary to know what the APPLIED yawing moment is (due to the engine-out moments), and to plot the applied moments on the same plot, in a similar form. It is possible to do this by choosing a gross weight to be used to calculate C L, and since standard bank angle will be 5°, the only remaining variable in CLsin φ is Ve. By choosing the appropriate values of Fn available versus Ve, a plot of Cna versus CLsinφ can be made which represents the applied yawing moments. If the weight chosen represents some standard minimum weight, and the available net thrust values represent the maximum allowable thrust scheduled in the AFM, the intersection point of the airframe curve and the engine curve will be the desired standard day values which can be used to calculate V MCA.
As an example, the following values of net thrust plus windmill drag have been extracted from a typical corporate jet engine spec. The data represent a maximum thrust engine, and have been corrected for ram drag and minimum accessory bleed and electrical load. The CLsin φ values are based on a gross weight of 9000 lbs.
Ve Fna Cna CLsin φ 70 2,846 0.079 0.204 90 2,798 0.047 0.123 110 2,764 0.031 0.082 130 2,737 0.022 0.059 150 2,710 0.016 0.044 Plotting both the airframe and engine yawing moment curves on the same graph looks like:
Airframe .048 Engine .124 0
0.01
0.02
0.03
0.04
0.05
0.06
0.07
0.08
0.09 0 0.05 0.1 0.15 0.2 0.25 CLsinφ Cna The intersection of the airframe and the engine curve shows values of:
Cna = 0.048 CLsin φ = 0.124 Since the engine Cn curve was based on W = 9000 lbs, the standard day value of V MCA can be determined from:
CLsin φ = 0.124 = 295 WV Se2 ⋅⋅ V .1108 9000
0.124e 2 = ⋅
VMCA = 89.7 KEAS If the airframe data is obtained from flight test, there are no assumptions or simplifications, and the value of V MCA derived from this method includes all the effects of bank angle, sideslip, adverse yaw, angle of attack, etc. Also, since the standard day value of Cna will always be less than the test value, no extrapolation is required, and there is no restriction on the value of standard day thrust that may be used.
Equation Method A single test day value of V MCA can also be corrected to standard conditions (using all the appropriate variables) without using this graphical method, provided either the slope of the Cna vs CLsin φ relationship is known (from wind tunnel, or analytical estimates), or one is willing to use a default (conservative) value. Thrust extrapolation using slope values not based on flight test is limited to 10% of the test day thrust. The following analysis shows the derivation of this single-point correction equation:
If the test day engine Cna curve was added to the previous plot, it would be possible to see how far, and in what direction, the correction from test to standard day was made. Assuming that a single value of V MCA was determined at 3000 feet at a weight of 9000 lbs, a test day engine Cna curve could be plotted using the same technique used for the standard day curve, except substituting the 3000 ft thrust values from the engine spec:
Ve Fna Cna CLsin φ 70 2,634 0.073 0.204 90 2,589 0.043 0.123 110 2,554 0.029 0.082 130 2,528 0.02 0.059 150 2,510 0.015 0.044