Change 1 114 3 For approach configurations, speeds from 1.23 V SR to VFE or VLE, as appropriate, and power from idle to go-around power/thrust; and
4 For landing configurations, speeds from V REF-5 knots to VFE or VLE, as appropriate, with power from idle to go-around power/thrust at speeds from VREF to VFE/VLE, and idle power at VREF-5 knots (to cover the landing flare).
b. Procedures . The test conditions should include each flap and landing gear configuration as described in paragraphs 27a(1) through 27a(4) at an altitude appropriate to each configuration.
(1) Basic Tests for Static Directional and Lateral Stability .
(a) Static Directional Stability. To check static directional stability with the airplane in the desired configuration and stabilized at the trim speed, the airplane is slowly yawed in both directions while maintaining the wings level with ailerons. When the rudder is released, the airplane should tend to return to straight flight.
(b) Static Lateral Stability. To check lateral stability with a particular configuration and trim speed, conduct steady, straight sideslips at the trim speed by maintaining the airplane heading with rudder and banking with aileron. When the ailerons are released, with the rudder held fixed, the low wing should tend to return to level. Initial bank angle should be appropriate to type; however, it is recommended that it should not be less than 10 degrees or that necessary to maintain the steady, straight sideslip with one-half rudder deflection, whichever occurs first. Roll control centering by the pilot should not be permitted during this evaluation.
The intent of this testing is to evaluate the short-term response of the airplane; therefore longterm effects, such as those due to spanwise fuel movement, need not be taken into account.
(2) Steady, Straight Sideslips. Steady, straight sideslips should be conducted in each direction to show that the aileron and rudder control movements and forces are substantially proportional to the angle of sideslip in a stable sense, and that the factor of proportionality is within the limits found necessary for safe operation. These tests should be conducted at progressively greater sideslip angles up to the sideslip angle appropriate to the operation of the airplane (see paragraph 27a(3)(b)) or the sideslip angle associated with one-half of the available rudder control input (as limited by a rudder control force of 180 pounds), whichever is greater.
(a) When determining the rudder and aileron control forces, the controls should be relaxed at each point to find the minimum force needed to maintain the control surface deflection. If excessive friction is present, the resulting low forces will indicate the airplane does not have acceptable stability characteristics.
(b) In lieu of conducting each of the separate qualitative tests described in paragraph 27b(1), the applicant may use recorded quantitative data showing aileron and rudder control force and position versus sideslip (left and right) to the appropriate limits in the steady heading sideslips conducted to show compliance with § 25.177(c). If the control force and position versus sideslip indicates positive dihedral effect and positive directional stability, compliance with § 25.177(a) and (b) will have been successfully demonstrated.
Change 1 115
(3) Full Rudder Sideslips .
(a) Rudder lock is that condition where the rudder over-balances aerodynamically and either deflects fully with no additional pilot input or does not tend to return to neutral when the pilot input is released. It is indicated by a reversal in the rudder control force as sideslip angle is increased. Full rudder sideslips are conducted to determine the rudder control forces and deflections out to sideslip angles associated with full rudder control input (or as limited by a rudder control force of 180 pounds) to investigate the potential for rudder lock and lack of directional stability.
(b) To check for positive directional stability and for the absence of rudder lock, conduct steady heading sideslips at increasing sideslip angles until obtaining full rudder control input or a rudder control force of 180 pounds. If full lateral control is reached before reaching the rudder control limit or 180 pounds of rudder control force, continue the test to the rudder limiting condition in a non-steady heading sideslip maneuver.
(4) Control Limits. The control limits approved for the airplane should not be exceeded when conducting the flight tests required by § 25.177.
(5) Flight Test Safety Concerns. In planning for and conducting the full rudder sideslips, items relevant to flight test safety should be considered, including:
(a) Inadvertent stalls,
(b) Effects of sideslip on stall protection systems,
(c) Actuation of stick pusher, including the effects of sideslip on angle-of-attack sensor vanes,
(d) Heavy buffet,
(e) Exceeding flap loads or other structural limits,
(f) Extreme bank angles,
(g) Propulsion system behavior (e.g., propeller stress, fuel and oil supply, and inlet stability),
(h) Minimum altitude for recovery,
(i) Resulting roll rates when the aileron limit is exceeded, and
(j) Position errors and effects on electronic or augmented flight control systems, especially when using the airplane’s production airspeed system.
Change 1 116
(k) Rudder loads, particularly those that may occur with dynamic rudder inputs.
28. Dynamic Stability - § 25.181.
a. Explanation .
(1) The dynamic stability tests described in this section should be conducted over the speed range of 1.13 VSR to VFE, VLE or VFC/MFC, as appropriate.
(2) Dynamic Longitudinal Stability .
(a) The short period oscillation is the first oscillation the pilot sees after disturbing the airplane from its trim condition with the pitch control (as opposed to the long period (phugoid)). Care should be taken that the control movement used to excite the motion is not too abrupt.
(b) Heavily damped means that the oscillation has decreased to 1/10 the initial amplitude within approximately two cycles after completion of the control input.
(c) Short period oscillations must be heavily damped, both with controls free and controls fixed.
(3) Dynamic Lateral-Directional Stability . The evaluation of the dynamic lateraldirectional stability should include any combined lateral-directional oscillation (“Dutch roll”) occurring over the speed range appropriate to the airplane configuration. This oscillation must be positively damped with controls free and must be controllable with normal use of the primary controls without requiring exceptional piloting skill.
b. Procedures .
(1) Dynamic Longitudinal Stability .
(a) The test for longitudinal dynamic stability is accomplished by a rapid movement or pulse of the longitudinal control in a nose up and nose down direction at a rate and degree necessary to obtain a short period pitch response from the airplane.
(b) Dynamic longitudinal stability should be checked at a sufficient number of points in each configuration to assure compliance at all operational speeds.
(2) Dynamic Lateral-Directional Stability .
(a) A typical test for lateral-directional dynamic stability is accomplished by a rudder doublet input at a rate and amplitude that will excite the lateral-directional response (“Dutch roll”). The control input should be in phase with the airplane’s oscillatory response.
(b) Dynamic lateral-directional stability should be checked under all conditions and configurations. If critical, special emphasis should be placed on adverse wing fuel loading conditions.
(3) Airplanes Equipped with Stability Augmentation Systems (SAS) . In the event a SAS is required for the airplane to show compliance with § 25.181(a) or (b), it must meet the requirements of §§ 25.671 and 25.672. Additionally:
(a) If the airplane is equipped with only one SAS (i.e., a single strand system), in accordance with § 25.672, compliance with the dynamic stability requirements of § 25.181(a) or
(b), as applicable, must be shown throughout the normal operating flight envelope to be certificated with the SAS operating, and in a reduced, practical operating flight envelope that will permit continued safe flight and landing with the SAS inoperative.
(b) If the airplane is equipped with more than one SAS, the resulting effects of SAS failure should be considered when determining whether or not the primary and any redundant SAS should be operating simultaneously for showing compliance with the dynamic stability requirements of § 25.181(a) or (b). If the primary and redundant SAS are dissimilar, the functional capability (i.e., control authority) of the redundant SAS should be considered with regard to restricting the operating envelope after failure of the primary SAS. At the applicant’s option, however, compliance with § 25.181(a) or (b) may still be demonstrated to a reduced flight envelope with no SAS operating as described in paragraph 28b(3)(a), above.
(c) Regardless of the SAS redundancy, the airplane should be safely controllable at the point of system failure or malfunction anywhere in the approved operating flight envelope of the airplane. Accordingly, it should be demonstrated that the airplane remains controllable during transition from the operating SAS to any redundant SAS, and during transition from anywhere in the normal operating envelope to the reduced practical operating envelope of § 25.672(c), if applicable. Airplane controllability should be demonstrated to meet the following levels as defined by the FAA HQRM. (The FAA HQRM is described in Appendix 6 to this AC.)
1 In the normal operating flight envelope with the SAS operating, the handling qualities should be “satisfactory” (SAT) as defined by the FAA HQRM.
2 At the point of SAS failure in the normal operating envelope, the airplane should be “controllable” (CON), as defined by the FAA HQRM, during the short term transitory period required to attain a speed and configuration that will permit compliance with paragraph 3, below.
118 3 During transition from the primary SAS to a redundant SAS, or from the normal operating envelope to a reduced, practical operating envelope (where applicable), the handling qualities should be “adequate” (ADQ) as defined by the HQRM.
4 In the reduced, practical operating flight envelope that will permit continued safe flight and landing, the handling qualities should be “satisfactory” (SAT) as defined by the HQRM.
Section 6. Stalls
29. Stall Testing.
a. The applicable Code of Federal Regulations (CFR) are as follows:
Section 25.21(c) Proof of Compliance
Section 25.103 Stall Speed
Section 25.143 Controllability and Maneuverability (General)
Section 25.201 Stall Demonstration
Section 25.203 Stall Characteristics
Section 25.207 Stall Warning
b. Explanation .
(1) The purpose of stall testing is threefold:
(a) To define the reference stall speeds and how they vary with weight, altitude, and airplane configuration.
(b) To demonstrate that handling qualities are adequate to allow a safe recovery from the highest angle-of-attack attainable in normal flight (stall characteristics).
(c) To determine that there is adequate pre-stall warning (either aerodynamic or artificial) to allow the pilot time to recover from any probable high angle-of-attack condition without inadvertently stalling the airplane.
(2) During this testing , the angle-of-attack should be increased at least to the point where the behavior of the airplane gives the pilot a clear and distinctive indication through the inherent flight characteristics or the characteristics resulting from the operation of a stall identification device (e.g., a stick pusher) that the airplane is stalled.
c. Stall Demonstration - § 25.201.
(1) The airplane is considered to be fully stalled when any one or a combination of the characteristics listed below occurs to give the pilot a clear and distinctive indication to cease any further increase in angle-of-attack, at which time recovery should be initiated using normal techniques.
(a) The pitch control reaches the aft stop and is held full aft for two seconds, or until the pitch attitude stops increasing, whichever occurs later. In the case of turning flight stalls, recovery may be initiated once the pitch control reaches the aft stop when accompanied by a rolling motion that is not immediately controllable (provided the rolling motion complies with § 25.203(c)).
(b) An uncommanded, distinctive and easily recognizable nose down pitch that cannot be readily arrested. This nose down pitch may be accompanied by a rolling motion that is not immediately controllable, provided that the rolling motion complies with § 25.203(b) or (c), as appropriate.
(c) The airplane demonstrates an unmistakable, inherent aerodynamic warning of a magnitude and severity that is a strong and effective deterrent to further speed reduction. This deterrent level of aerodynamic warning (i.e., buffet) should be of a much greater magnitude than the initial buffet ordinarily associated with stall warning. An example is a large transport airplane that exhibits “deterrent buffet” with flaps up and is characterized by an intensity that inhibits reading cockpit instruments and would require a strong determined effort by the pilot to increase the angle-of-attack any further.
(d) The activation point of a stall identification device that provides one of the characteristics listed above. Paragraph 228 of this AC presents guidance material for demonstrating compliance with the regulatory requirements of part 25 for stall identification systems.
(2) It should be recognized that the point at which the airplane is considered stalled may vary, depending on the airplane configuration (e.g., flaps, gear, center of gravity, and gross weight). In any case, the angle-of-attack should be increased until one or more of these characteristics is reached for all likely combinations of variables.
d. Stall Speeds .
(1) Background .
(a) Since many of the regulations pertaining to performance and handling qualities specify trim speeds and other variables that are functions of stall speeds, it is desirable to accomplish the stall speed testing early in the program, so the data are available for subsequent testing. Because of this interrelationship between the stall speeds and other critical performance parameters, it is essential that accurate measurement methods be used. Most standard airplane pitot-static systems are unacceptable for stall speed determination. These tests require the use of properly calibrated instruments and usually require a separate test airspeed system.
(b) Prior to amendment 25-108, the stall speed defined in § 25.103 was the minimum speed attained in the stalling maneuver. For many high speed swept wing transport category airplanes, the resulting stall speed often occurs at a load factor normal to the flight path considerably less than one, which leads to inconsistent and unrepeatable reference stall speeds.
Pilot technique can also significantly influence the rate and magnitude of any nose down pitch occurring at the stall, thereby contributing to inconsistencies in the determination of the minimum speed obtained in the stalling maneuver. Since part 25 defines operating speeds as multiples of the stall speed, the resulting operating speed margins to stall may not be representative of the actual lift margin available (i.e., the margin to the speed at which wing lift alone can support the weight of the airplane in 1-g flight). To ensure that operating speed and maneuvering margins are directly related to wing lift margin, amendment 25-108 redefined the reference stall speed as the 1-g stall speed, which is the speed at which the wing is generating maximum usable lift in a 1-g flight condition.
(c) Since the 1-g stall speed is generally higher than the minimum speed obtained in the stalling maneuver, retaining the existing multiplying factors for determining the minimum operating speeds would have resulted in higher minimum operating speeds. To retain approximately equivalent operating speeds that have been found acceptable, the multiplying factors applied to the reference stall speed were reduced by approximately six percent.
(d) Although the reduced multiplying factors were intended to result in roughly equivalent operating speeds, there is one class of airplanes for which a significantly lower operating speed would be obtained. Airplanes equipped with a device that abruptly pushes the
121 nose down (e.g., a stick pusher) near the angle-of-attack for maximum lift would be operated at speeds and angles-of-attack closer to the pusher activation point than has been experienced in operational service. Therefore, to maintain equivalency in operating speeds for these airplanes, a supplementary margin has been established such that VSR must not be less than the greater of 2 knots or 2 percent above the speed at which the device activates (§ 25.103(d)). In addition, see paragraph 228 of this AC for guidance material regarding the means of compliance criteria for the design and function of such systems.
(2) Configuration .
(a) Stall speeds should be determined for all aerodynamic configurations to be certificated for use in the takeoff, en route, approach, and landing configurations.
(b) The center of gravity positions to be used should be those that result in the highest stall speeds for each weight (forward c.g. in most cases).
(c) Sufficient testing should be conducted to determine the effects of weight on stall speed. Altitude effects (compressibility, Reynolds Number) may also be considered if credit for variations in these parameters is sought by the applicant. If stall speeds are not to be defined as a function of altitude, then all stall speed testing should be conducted at a nominal altitude no lower than 1,500 ft. above the maximum approved takeoff and landing altitude. (See paragraph 29d(5)(g).)
(3) Procedures .
(a) The airplane should be trimmed for hands-off flight at a speed 13 percent to 30 percent above the anticipated VSR, with the engines at idle and the airplane in the configuration for which the stall speed is being determined. Then, using only the primary longitudinal control for speed reduction, maintain a constant deceleration (entry rate) until the airplane is stalled, as defined in § 25.201(d) and paragraph 29c(1) of this AC. Following the stall, engine thrust may be used as desired to expedite recovery.
(b) A sufficient number of stalls (normally four to eight) should be accomplished at each critical combination of weight, altitude, c.g., and external configuration. The intent is to obtain enough data to define the stall speed at an entry rate of 1.0 knot/second.
(c) During the stall speed testing, the stall characteristics of the airplane must also satisfy the requirements of § 25.203(a) and (b).
(d) For airplanes that have stall identification devices for which the angle-ofattack for activation is biased by angle-of-attack rate, some additional considerations are necessary. The stall speeds are normalized against an average airspeed deceleration rate, as described in paragraph 29d(5)(e). However, stall identification systems generally activate at a specific angle-of-attack, biased by an instantaneous angle-of-attack rate. Therefore, longitudinal control manipulation by the pilot during the stall maneuver, close to the stall identification system activation point, can advance or delay its activation without appreciably affecting the
122 average stall entry airspeed rate. To minimize scatter in the stall speed versus entry rate data, the pilot should attempt to maintain a stable angle-of-attack rate or pitch rate (not necessarily a fixed airspeed deceleration rate), until the stall identification system activates. The resulting timehistory of angle-of-attack data should be smooth and without discontinuities. A cross plot of airspeed deceleration rate, as defined in paragraph 29d(5)(e), versus angle-of-attack rate for all related test points, will show the general trend of this relationship for each flap setting. Any points that do not follow this general trend should not be used in establishing the stall speed.
(4) Thrust Effects on Stall Speed .
(a) Stall speeds are typically determined with the thrust levers at idle; however, it is necessary to verify by test or analysis that engine idle thrust does not result in appreciably lower stall speeds than would be obtained at zero thrust. Prior to amendment 25-108, a negative idle thrust at the stall, which slightly increases stall speeds, was considered acceptable, but applicants were not required to base stall speeds on idle thrust. With the adoption of amendment 25-108, it became a requirement to base stall speeds on idle thrust, except where that thrust level results in a significant decrease in stall speeds. If idle thrust results in a significant decrease in stall speeds, then stall speeds cannot be based on more than zero thrust.
(b) To determine whether thrust effects on stall speed are significant, at least three stalls should be conducted at one flap setting, with thrust set to approximately the value required to maintain level flight at 1.5 VSR in the selected configuration.
(c) These data may then be extrapolated to a zero thrust condition to determine the effects of idle thrust on stall speeds (see figure 29-1). If the difference between idle thrust and zero thrust stall speed is 0.5 knots or less, the effect may be considered insignificant.
Figure 29-1. Thrust Effect On Stall Speed
(d) The effects of engine power on stall speeds for a turbopropeller airplane can be evaluated in a similar manner. Stall speed flight tests should be accomplished with engines idling and the propellers in the takeoff position. Engine torque, engine r.p.m., and estimated propeller efficiency can be used to predict the thrust associated with this configuration.
(5) Data Reduction and Presentation . The following is an example of how the data obtained during the stall speed testing may be reduced to standard conditions. Other methods may be found acceptable.
(a) Record the indicated airspeed from the flight test airspeed system throughout the stall, and correct these values to equivalent airspeed. Also record load factor normal to the flight path. Typically, the load factor data would be obtained from a sufficient number of accelerometers capable of resolving the flight path load factor. It may be possible to obtain acceptable data using one accelerometer aligned along the expected 1-g stall pitch angle. More likely, it will take at least two accelerometers, one aligned along the fuselage longitudinal axis and one aligned at 90 degrees to that axis, as well as a means to determine the angle between the flight path and the fuselage longitudinal axis.
(b) Calculate the airplane lift coefficient (C L) from the equation given below and plot it as a time history throughout the stall maneuver.
C L = nWqSzw = SV Wne zw237.295
ENGINE THRUST
STALL SPEED
ZERO THRUST THRUST FOR 1.5 VSR
IN LEVEL FLIGHT IDLE THRUST
THRUST EFFECT ON
STALL SPEED
Where: n zw = airplane load factor normal to the flight path W = airplane test weight - lbs.
q = dynamic pressure - lbs./ft. 2 S = reference wing area - ft. 2 Ve = knots equivalent airspeed.
(c) The maximum lift coefficient is defined as the maximum value of C L achieved during the stall test. Where the time history plot of CL exhibits multiple peak values, corresponds to the first maximum. There should also typically be a noticeable break in a plot of the load factor normal to the flight path near the point at which is reached. The analysis to determine should disregard any transient or dynamic increases in recorded load factor, such as might be generated by abrupt control inputs that do not reflect the lift capability of the airplane. The load factor normal to the flight path should be maintained at nominally 1.0 until is reached. (See figure 29-2.)
Figure 29-2. CLMAX and Load Factor
CLMAX (CLMAX) CLMAX CLMAX CLMAX
(d) Correct the obtained for each stall, if necessary, from the test c.g.
position to the targeted c.g. position, and for any thrust effects, using the equation:
C LMAX = C LMAX (test c..g. position)[1 + (MAC/lt)(CGstd - CGtest)] - CLT
Where: MAC = Wing mean aerodynamic chord length -` inches.
l t = Effective tail length, measured between the wing 25 percent MAC and the stabilizer 25 percent MAC - inches.
C G std = C.G. position resulting in the highest value of reference stall speed (normally the forward c.g. limit at the pertinent weight)
- percent MAC/100
C G test = Actual test c.g. position - percent MAC/100
CLT = Change in C L due to engine thrust (if effect of idle thrust is greater than 0.5 knots in stall speed).
(e) Determine the stall entry rate, which is defined as the slope of a straight line connecting the stall speed and an airspeed 10 percent above the stall speed, for each stall test.
Because is relatively insensitive to stall entry rate, a rigorous investigation of entry rate effects should not be necessary. Test data should bracket a 1.0 knot/second entry rate such that the value of corresponding to an entry rate of 1.0 knot/second can be determined. (See figure 29-3.)