AC 25-7C

Appendix 5 A5-9 Figure 6. Flaps DOWN Flight Envelopes

V

DET

or or -150

*

*

0.8

0.5

0.0

or IB

 ENG INOP

or

or

1.6

1.3

LIM DYN

 LIM

 LIM DYN

2.0 LIM   LIM

or 

V

ALTITUDE AOA ROLL BETA BUFFET V PITCH V PITCH    VFE VFE + V sw CLMAX

CLMAX LIM STDY LIM STDY VSW j or ±670 ±500 ±400 V2+X X ±400 LAND NFE NFE OFE LFE +300 NZ NZ Normal Flight Envelope Operational Flight Envelope Limit Flight Envelope NFE LFE OFE LIM VFE + V TO & L +3000 TO & L +3000 ±300

TAKEOFF VRFE * * V *Rudder appropriate to task/operation/ or rudder lock evaluation

Appendix 5 A5-10

f. Figure 7 presents the method for combining the various flight condition parameter probabilities to determine the minimum acceptable handling qualities rating for each combination of these parameters. This method is shown graphically in Figure 8.

Figure 7. Combining Values

A. Analyze Failures/Determine Flight Control System Failure Probability (Xc)

-- Predicted failure rates/check failure co-dependence -- Equipment Inoperative Dispatch under MEL -- Service Difficulty Records (continuing airworthiness)

B. Determine Flight Envelope Probabilities (Xe) and Atmospheric Probabilities (Xa) for the flight condition

C. Modify the flight envelope probability if inter-related with the atmospheric condition (See Figure 4, Section D.)

D. Repeat process to identify all cases where Xc * Xa * Xe   10 -9

E. Determine: “Flight Condition” (Xc * Xe)

-- PROBABLE FLIGHT CONDITION: 10-5  (Xc * Xe) < 0 -- IMPROBABLE FLIGHT CONDITION: 10-9  (Xc * Xe) < 10-5

Appendix 5 A5-11 Figure 8. Probability Guidelines to Determine HQ Requirements

0 -1 -2 -3 -4 -5 -6 -7 -8 -9

PROBABLE FAIL IMPROBABLE FAIL EFCS FAILURE RATE -10Xc Xc= -9 -8 -7 -6 -5 -4 -3 -2 -1 0 EFCS *Based on (Xe+Xc) Xe Xc Xa IMPROB FLT CONDPROB FLT COND* PROBABILITY -10x

FLIGHT ENVELOPE &

ATMOSPHERIC DISTURBANCE **Gust intensity variations may be expected with altitude and between u, v, w axes

OFE LFENFE

MOD SEVERELIGHT 3 5 8 14 20GUST** fps SHEAR XWIND 10 18 25 33 45 kts

SMALL RAMP INCR & JFK DEN MSY DFW (TO&L) ug FLIGHT

ENVELOPE

ATMOSPHERIC

DISTURBANCE

Appendix 5 A5-12

g. Figure 9 shows the minimum acceptable FAA handling qualities rating for a given flight condition, defined as a combination of the flight envelope conditions and the level of atmospheric disturbance, relative to the probability of the failure condition being evaluated.

Figure 9 is not meant to imply that every atmospheric disturbance level and every flight envelope combination must be tested. It simply shows the minimum acceptable handling qualities rating for a handling qualities task conducted in a specific environmental state (i.e., atmospheric disturbance level), in a specific segment of the flight envelope, and in a specific system failure state.

Figure 9. Minimum HQ Requirements

ATMOSPHERIC DISTURBANCE (Xa)

FLIGHT LIGHT MODERATE SEVERE CONDITION FLIGHT ENVELOPE (Xe) (Xc*Xe) NFE OFE LFE NFE OFE LFE NFE OFE LFE

PROBABLE

CONDITION S S A A C C C C C

IMPROBABLE

CONDITION A A C C C C

S=SATISFACTORY A=ADEQUATE C=CONTROLLABLE

NFE – Normal Flight Envelope

OFE – Operational Flight Envelope

LFE – Limit Flight Envelope

Appendix 6 A6-1

Appendix 6

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 7.

Theoretical Basis

Given the static lateral/directional equations of motion for straight line, unaccelerated flight:

Fy  Cy   a yaC    Cy rr  = CL · sin 

Mx  Cl   Cl aa   Cl rr    

Mz  Cn   Cn aa   Cn rr  Cna  

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

Appendix 6 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, r is a constant, and the system of equations can be resolved to an identity which shows that Cna is a linear function of CLsin .

 Cna = A·CL  sin + 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 VMCA can be corrected to standard conditions by the relationship:

= · for Turbojets = · for Propeller driven where:

THPs = Maximum AFM scheduled Brake/Shaft horsepower multiplied by standard day propeller efficiency.

THPt = Test day brake/shaft horsepower where VMCA was achieved multiplied by test day propeller efficiency.

= Atmospheric density ratio at standard conditions.

= Atmospheric density ratio at test conditions.

Windmilling shaft horsepower is not considered, because the current 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 VMCA due to variations in bank angle, weight, sideslip angle, or adverse yaw, its use is limited to corrections of 5 percent or less in asymmetric net thrust or power.

A6-2

Appendix 6 A6-3 For corrections beyond 5 percent, the relationship shown in equation (4) should be used, and enough flight test data should be obtained to define the correlation between Cna and CLsin.

Graphical Method

In theory these data could be obtained by varying any combination of asymmetric power or thrust, airspeed, weight, and bank angle that would provide a representative variable set.

However, since VMCA 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 power or thrust will cause VMCA 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, VMCA tests are normally done at the lightest possible weight, at the maximum allowable asymmetric power or 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 VMCA 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 power or thrust, testing at three bank angles is required for the definition of the Cna vs CL sin relationship. These data should be obtained by shutting down the critical engine (normally the left), setting maximum allowable power or 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.

Appendix 6 A6-4 Data obtained using this method with a typical business jet and a large jet transport are shown in the following two graphs:

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 CL, and since standard bank angle will be 5°, the only remaining variable in CLsin is Ve. By choosing the appropriate values of Fn available versus Ve, a plot of Cna versus CLsin can be made which represents the applied yawing moments. If the weight chosen represents some standard minimum weight, and the available net power or thrust values represent the maximum

Appendix 6 A6-5 allowable power or 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 VMCA.

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 values are based on a gross weight of 9000 lbs.

Ve Fna Cna CLsin 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 CLsin Cna

The intersection of the airframe and the engine curve shows values of:

Cna = 0.048 CLsin = 0.124

Since the engine Cn curve was based on W = 9000 lbs, the standard day value of VMCA can be determined from:

CLsin = 0.124 = 295 W V Se2  

Appendix 6 A6-6

0.124 90000.1108V 2e 

VMCA = 89.7 KEAS

If the airframe data is obtained from flight test, there are no assumptions or simplifications, and the value of VMCA 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 power or thrust that may be used.

Equation Method

A single test day value of VMCA 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 relationship is known (from wind tunnel, or analytical estimates), or one is willing to use a default (conservative) value. Power or thrust extrapolation using slope values not based on flight test is limited to 10 percent of the test day power or 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 VMCA 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 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

Appendix 6 The following Cna vs CLsin plot shows the airframe curve, the standard day engine curve, and the test day curve:

From the airframe/engine curve intersections:

Cnt = 0.052 C ns = 0.048 CLsint = 0.147 C Lsins = 0.124 = 82.4 KEAS = 89.7 KEAS This is a thrust correction of approximately 8 percent. If the constant Cn method had been used, the test value of .052 would have applied, and the corresponding would have been 85.7 KEAS, an error of 4 knots (5 percent) in the non-conservative direction.

Noting that the correction from test day to standard day is along the Cn vs CLsin curve, which is a straight line, and denoting the slope of the airframe curve as K the intersection of the standard day thrust line with the airframe curve as Cns and CLsins, and the intersection of the test day thrust line with the airframe curve as Cnt and CLsint, the following equation can be derived:

Given the point-slope form of a straight line,

Appendix 6 A6-8 Cns - Cnt = K  (CLsins - CLsint ) Cns - K  CLsins = Cnt - K  CLsint F l q S b na e s s  

- K   W sin q S s s s    F l q S b na e t t    K  W sin q S t t t   Substituting q = V 295 e 2 and then multiplying through by V S 295 e 2 S  Fnas  l b e - K  Ws sins = V V e 2 e 2 s t  F lb K W sinnat e t t      and finally, VMCAs = VMCAt  F lb K W sin F lb K W sin na e s s na e t t s t                           1 2 for Turbojets VMCAs = VMCAt  326 SHPV lb K W sin 326 SHPV lb K W sin S S S MCA e s s t t t MCAt e t t s                                  1 2 for Turboprops In the simplified form of the lateral directional equations, K is which is the directional static margin. The value of K typically varies from approximately 0.14 to 0.19, depending on the lateral/directional characteristics of the airplane being tested. As with the graphical method, if K is determined by flight test, power or thrust corrections to VMCA are based on an interpolation of flight test defined airframe capability, and there is no limit on the amount of the power or thrust correction to VMCA. A somewhat conservative default value of

0.20 for K may be used if flight test data is not available; however, in this case, any power or thrust extrapolation is limited to 10 percent of the test day power or thrust. To assure that corrections for bank angle and weight do not result in standard day VMCA values at or below stall speed, the corrections made by either the graphical or equation method should not result in a VMCA which is based on a CLsin that is greater than CLMAX sin5.

C Cn y 