Section 4. TRIM
24. [RESERVED]
Section 5. STABILITY
25. GENERAL - § 25.171. [RESERVED]
26. STATIC LONGITUDINAL STABILITY AND DEMONSTRATION OF STATIC LONGITUDINAL STABILITY - §§ 25.173 AND 25.175.
a. Explanation.
(1) Section 25.173 - Static Longitudinal Stability.
(i) Compliance with the general requirements of § 25.173 is determined from a demonstration of static longitudinal stability under the conditions specified in § 25.175.
(ii) The requirement is to have a pull force to obtain and maintain speeds lower than trim speed, and a push force to obtain and maintain speeds higher than trim speed. There may be no force reversal at any speed that can be obtained, except lower than the minimum for steady, unstalled flight or, higher than the landing gear or wing flap operating limit speed or VFC/MFC, whichever is appropriate for the test configuration. The required trim speeds are specified in § 25.175.
(iii) When the control force is slowly released from any speed within the required test speed range, the airspeed must return to within 10 percent of the original trim speed in the climb, approach, and landing conditions, and return to within 7.5 percent of the trim speed in the cruising condition specified in § 25.175 (free return).
(iv) The average gradient of the stick force versus speed curves for each test configuration may not be less than one pound for each 6 knots for the appropriate speed ranges specified in § 25.175. Therefore, after each curve is drawn, draw a straight line from the intersection of the curve and the required maximum speed to the trim point. Then draw a straight line from the intersection of the curve and the required minimum speed to the trim point. The slope of these lines must be at least one pound for each 6 knots. The local slope of the curve must remain stable for this range.
NOTE: Due to different design features of individual airplanes, there may be cases where the local slope gradient deviates somewhat from that specified by § 25.173. When this occurs, an investigation should be performed to determine if a finding of equivalent safety can be made based on pilot evaluation.
(2) Section 25.175, Demonstration of Static Longitudinal Stability, specifically defines the flight conditions, airplane configurations, trim speed, test speed ranges, and thrust settings to be used in demonstrating compliance with the longitudinal stability requirements.
b. Procedures.
(1) For the demonstration of static longitudinal stability, the airplane should be trimmed in smooth air at the conditions required by the regulation. Aft c.g. loadings are
generally most critical. After obtaining trim speed, apply a light pull force and stabilize at a slower speed. Continue this process in acceptable increments, depending on the speed spread being investigated, until reaching the minimum speed for steady, unstalled flight or the minimum required as appropriate for the configuration. A continuous pull force from the trim speed is required on each series of test points to eliminate hysteresis effects. At the end of the required speed range, the force should be gradually relaxed to allow the airplane to slowly return toward the trim speed and zero stick force. Depending on the amount of friction in the control system, the eventual speed at which the airplane stabilizes will normally be less than the original trim speed. The new speed, called the free return speed, must meet the requirements of § 25.173.
(2) Starting again at the trim speed, push forces should be gradually applied and gradually relaxed in the same manner as described in paragraph (1), above.
(3) The above techniques result in several problems in practice. One effect of changing airspeed is a change of altitude, with a corresponding change in Mach number and thrust or power output. Consequently, a reasonably small altitude band, limited to +3,000 ft., should be used for the complete maneuver. If the altitude band is exceeded, regain the original trim altitude by changing the power setting and flap and gear position, but without changing the trim setting. Then continue the push or pull maneuver in the original configuration. Testing somewhat beyond the required speed limits in each direction assures that the resulting data at least extends to the required speed ranges. It will also be noted in testing that while holding force constant at each data point, the airspeed and instantaneous vertical speed vary in a cyclic manner. This is due to the long period (phugoid) oscillation. Care should be exercised in eventing the data point, since it may be biased by this phugoid oscillation. Averaging these oscillating speeds at each data point is an acceptable method of eliminating this effect.
Extremely smooth air improves the quality of the test data, with critical areas requiring the best of smooth air. In-bay and cross-bay wing fuel shift is another problem experienced in some airplanes. In-bay fuel shift occurs rapidly with pitch angle; therefore, consideration should be given to testing with fuel loadings that provide the maximum shift since it is generally destabilizing. Slower, cross-bay fuel shift, or burn from an aft tank, can influence the measured stability but usually only because of the time required to obtain the data points. This testing induced instability should be removed from the data before evaluation of the required slope.
(4) The resulting pilot longitudinal force test points should be plotted versus airspeed to show the positive stable gradient of static longitudinal stability and that there are no “local” reversals in the stick force vs. airspeed relationship over the range of airspeeds tested. This plot should also show the initial trim point and the two return-to-trim points to evaluate the return-totrim characteristics (see Figure 26-1).
FIGURE 26-1. LONGITUDINAL STATIC STABILITY
27. STATIC DIRECTIONAL AND LATERAL STABILITY - § 25.177.
a. Explanation.
(l) Static Directional Stability. Positive static directional stability is defined as the tendency to recover from a skid with the rudder free. Prior to Amendment 25-72, a separate demonstration of positive static directional stability was required by § 25.177(a) for any landing gear and flap position and symmetrical power condition at speeds from 1.2VS1 up to VFE, VLE, or VFC/MFC, as appropriate for the airplane configuration.
(2) Static Lateral Stability. Positive static lateral stability is defined as the tendency to raise the low wing in a sideslip with the aileron controls free. Prior to Amendment 25-72, a separate demonstration was required by § 25.177(b) to show that static lateral stability was not negative in any landing gear and flap position and symmetrical power condition at speeds from
1.2VS1 to VFE, VLE, or VMO/MMO, as appropriate for the airplane configuration. At speeds from VMO/MMO to VFC/MFC, negative static lateral stability was permitted by § 25.177(b), providing the divergence was:
(i) Gradual;
(ii) Easily recognizable by the pilot; and
(iii) Easily controllable by the pilot.
(3) Steady Straight Sideslips.
(i) Section 25.177(c) requires, in steady, straight sideslips to sideslip angles appropriate to the operation of the airplane, that the aileron and rudder control movements and
forces be proportional to the angle of sideslip and must lie between limits necessary for safe operation. Experience has shown that 15 degrees is an appropriate sideslip angle for transport category airplanes. Section 25.177(c) states that compliance with these steady, straight sideslip criteria "must be demonstrated for all landing gear and flap positions and symmetrical power conditions at speeds from 1.2VS1 to VFE, VLE, or VFC/MFC, as appropriate."
(ii) At sideslip angles greater than those appropriate for normal operation of the airplane, up to the sideslip angle at which full rudder travel is used or a rudder pedal force of 180 pounds is obtained, the rudder pedal forces may not reverse and increased rudder deflection must produce increased angles of sideslip. The goals of this higher-than-normal sideslip angle test are to show that at full rudder, or at maximum expected pilot effort: (1) the rudder pedal force does not reverse, and (2) increased rudder deflection produces increased angles of sideslip, thus demonstrating adequate directional stability for maneuvers involving large, dynamic rudder inputs, such as engine failure during takeoff just after V1. Consequently, it is permissible to continue this test beyond full lateral control input.
(iii) Section 25.177(d) (post-Amendment 25-72) stipulates that increased rudder deflection must continue to increase the angle of sideslip at speeds between VMO/MMO and VFC/MFC; but the dihedral effect, as exhibited by aileron deflection opposite the corresponding rudder input, may become negative in this speed range, provided the divergence is:
(A) Gradual;
(B) Easily recognized by the pilot; and (C) Easily controlled by the pilot.
b. Procedures. The test conditions should include each flap and landing gear configuration as described in paragraphs 27a(1), (2) or (3), at both low altitude and the maximum altitude appropriate to each configuration.
(1) Basic Test.
(i) 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.
(ii) 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.
(iii) Absence of Rudder Lock.
(A) Rudder lock is that condition where the rudder over-balances aerodynamically and deflects fully with no additional pilot input.
(B) To check for the absence of rudder lock with a particular configuration and trim speed, conduct steady, straight sideslips (unaccelerated forward slips) while maintaining a desired airplane track.
(C) Aileron and rudder control movements must remain in harmony, and forces must increase in proportion with sideslip angle, up to the limits found necessary for safe operation. At sideslip angles greater than those appropriate for normal operation of the airplane, up to the sideslip angle at which full rudder travel is used or a rudder pedal force of 180 pounds is obtained, the rudder pedal forces may not reverse; and increased rudder pedal deflection must produce increased angles of sideslip. The purpose of the second part of this test, at greater than normal sideslip angles, is to show that no rudder lock tendencies exist up to the rudder deflection limits as may be encountered during a dynamic rudder input to maintain directional control.
(2) Alternative Test. In lieu of conducting each of the qualitative tests described in paragraph 27b(1), the applicant may obtain recorded quantitative data showing aileron and rudder force and position versus sideslip (left and right) to the appropriate limits in constant heading sideslips.
(i) When determining the rudder and aileron forces, the controls should be relaxed at each point to find the minimum force required to hold the control surface deflection. If excessive friction is present, the resulting low forces will indicate the airplane does not have acceptable stability characteristics.
(ii) If the force and position versus sideslip indicates positive dihedral effect, positive directional stability and no rudder lock, compliance with § 25.177 has been successfully demonstrated.
(3) The control limits approved for the airplane should not be exceeded when conducting the flight tests required by § 25.177.
28. DYNAMIC STABILITY - § 25.181.
a. Explanation.
(1) The dynamic stability tests described in this section should be conducted over the speed range of 1.2VS to VFE, VLE or VFC/MFC, as appropriate.
(2) Dynamic Longitudinal Stability.
(i) 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.
(ii) Heavily damped means that the oscillation has decreased to 1/10 the initial amplitude within approximately two cycles after completion of the control input.
(iii) 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.
(i) 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.
(ii) Dynamic longitudinal stability must be checked at a sufficient number of points in each configuration to assure compliance at all operational speeds.
(2) Dynamic Lateral-Directional Stability.
(i) 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.
(ii) Dynamic lateral-directional stability must 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:
(i) If the airplane is equipped with only one SAS (i.e., a single strand system), 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.
(ii) 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)(i), above.
(iii) Regardless of the SAS redundancy, the airplane must 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 Handling Qualities Rating System (HQRS). (The FAA HQRS is described in Appendix 7 to this AC.)
(A) In the normal operating flight envelope with the SAS operating, the handling qualities should be "satisfactory" (SAT) as defined by the FAA HQRS.
(B) At the point of SAS failure in the normal operating envelope, the airplane should be "controllable" (CON), as defined by the FAA HQRS, during the short term transitory period required to attain a speed and configuration that will permit compliance with paragraph (C), below.
(C) 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 HQRS.
(D) 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 HQRS.
Section 6. STALLS
29. STALL TESTING .
a. The applicable Federal Aviation Regulations (FAR) are as follows:
Section 25.21(c) Proof of Compliance Section 25.103 Stalling 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:
(i) To define the minimum inflight airspeeds and how they vary with weight, altitude, and airplane configuration (stall speeds).
(ii) To demonstrate that handling qualities are adequate to allow a safe recovery from the highest angle of attack attainable in normal flight (stall characteristics).
(iii) To determine that there is adequate pr e-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 following two conditions are satisfied:
(i) Attainment of an angle of attack measurably greater than that for maximum lift, except when the stall is defined by a stall identification device (e.g., stick pusher).
(ii) Clear indication to the pilot through the inherent flight characteristics or stall identification device (e.g., stick pusher) that the airplane is stalled.
(3) 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 that he should stop any further increase in angle of attack, at which time recovery should be initiated using normal techniques. The stall speed is defined as the minimum speed reached during the maneuver, except for those airplanes that require stall identification devices (see paragraph 29b(3)(iv), below).
(i) 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)).
(ii) 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.
(iii) 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) must 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.
(iv) The activation point of a stall identification device that is a strong and effective deterrent to further speed reduction. Paragraph 228 of this AC presents guidance material for demonstrating compliance of stall identification systems with the regulatory requirements of Part 25 of the FAR.
NOTE: The regulatory requirements of § 25.201(d) define the stalled condition as occurring when the airplane has reached an angle of attack measurably greater than that for maximum lift. The related stall speed has been defined as the minimum speed obtained in the stalling maneuver (V S, also referred to as V S MIN). This VS MIN definition has led to stall speeds sometimes being defined at excessively low load factors, with pilot objectivity and flying technique having considerable influence over the results. The subsequent effect on operating speeds can then lead to inconsistencies between similarly configured airplane designs and result in inadequate maneuver margins.
Consequently, applicants are encouraged to use 1-g stall speeds that reflect the ability of the wing-generated lift to support the weight of the airplane in 1-g flight. These 1-g stall speeds may be used as the reference datum for regulatory compliance through an equivalent safety finding. The FAA is in the process of amending the stall speed requirements to incorporate the 1-g stall speed methodology (ref. NPRM 95-17). Upon adoption of the final 1-g stall speed rule, further revisions to the guidance material in this AC will be proposed. Details of the 1-g stall speed regulatory equivalencies and the equivalent safety finding process are contained in Appendix 5.
(4) 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 must be increased until one or more of these characteristics is reached for all likely combinations of variables.
c. Stall Speeds .
(1) Background . 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 have not been found to be acceptable for stall speed determination.
These tests require the use of properly calibrated instruments and usually require a separate test airspeed system, such as a trailing bomb, a trailing cone, or an acceptable nose or wing boom.
(2) Configuration .
(i) Stall speeds should be determined for all aerodynamic configurations to be certificated for use in the takeoff, en route, approach, and landing configurations.
(ii) 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).
(iii) Sufficient testing should be c onducted 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 29c(5)(vii) and Figures 29 -3 and 29-4.)
(3) Procedures .
(i) The airplane should be trimmed for hands-off flight at a speed 20 percent to 40 percent above the anticipated stall speed, with the engines at idle and the airplane in the prescribed configuration for which the stall speed is being determined. Then, using only the primary longitudinal control, a constant deceleration (entry rate) is maintained until one of the previously defined points that define the stall is reached. Following the stall, engine thrust may be utilized, as desired, to expedite recovery.
(ii) A sufficient number of stalls (normally six) should be accomplished at each critical combination of weight, c.g., and external configuration, varying the entry rate from approximately 0.5 knots/second to 1.5 knots/second. The intent is to obtain enough data to define the stall speed at an entry rate of 1.0 knot/second.
(See Figure 29-2.)
(iii) During the stall speed testing, the stall characteristics of the airplane must also satisfy the requirements of §§ 25.203(a) and (b).
(iv) For airplanes that have stall identification devices for which the angle of attack 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 29c(5). 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 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 time-history of angle of attack data should be smooth and without discontinuities. A cross plot of airspeed deceleration rate, as defined in paragraph 29c(5), 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 .
(i) Stall speeds are normally defined with the thrust levers at idle; however, it is necessary to verify by test or analysis that engine idle thrust does not affect stall speeds to an extent that they are appreciably lower than would be experienced at zero thrust. Negative thrust at the stall, which slightly increases stall speeds, is acceptable.
(ii) 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 1.6 V S in the selected configuration.
(iii) These data may then be extrapolated to a zero thrust condition t o eliminate the effects of idle thrust on stall speeds (see Figure 29-5). If the difference between idle thrust and zero thrust stall speed is 0.5 knots or less, the effect may be considered insignificant.
(iv) The effects of engine power on stall spe eds for a turbopropeller airplane can be evaluated in a similar manner. Stall speed flight tests should be accomplished with throttles closed 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 have been found acceptable.
(i) Indicated airspeed from the flight test airspeed system is recorded throughout the stall, and these values are corrected to equivalent airspeed.
(ii) Stall entry rate is then defined as the slope of a straight line connecting stall speed and an airspeed 10 percent above the stall speed (see Figure 29-1).
Entry Rate = (VS-1.1 VS)/Time Difference
(iii) The airplane lift coefficient, C L, is then calculated for each test stall speed using the equation:
CL = W/qS = 295.37(W)/(V S(e)2S) Where: W = airplane test weight - lbs.
q = dynamic pressure - lbs./ft. 2 S = reference wing area - ft. 2 VS(e) = test stall speed corrected to knots equivalent airspeed.
(iv) The CL obtained for each stall is then corrected to the targeted c.g. position using the equation:
CLCG = CL[1+(MAC/lt)(CGstd -CGtest )] - ∆CLT Where: MAC = Wing mean aerodynamic chord length - inches.
lt = Effective tail length, measured between the wing 2 5 percent MAC and the horizontal stabilizer 25 percent MAC - inches.
CGstd = Forward c.g. limit at the pertinent weight - percent MAC/100 CGtest = Actual test c.g. position - percent MAC/100 ∆CLT = Change in CL due to engine thrust (if significant).
(v) For each targeted c.g. and external configuration, a plot of C LCG versus entry rate is constructed. The final stall C L is selected at an entry rate of 1.0 knot/second (see Figure 29-2).
(vi) For each final stall C L, weight, wing flap, and external configuration, a plot of stall C L versus weight is constructed (see Figure 29-3.) An initial negative slope of this plot may be caused by several factors:
(A) A decrease in CLMAX due to increasing Mach number (which increases as the stall speed goes up with weight);
(B) The fact that CLMAX is proportional to the rate of change of angle of attack, whereas the data are plotted at fixed airspeed bleed rate; and (C) Minor adverse aeroelastic effects on the wings and high lift devices as weight (and therefore speed) increases. The inflection at the right end is caused typically by a less forward c.g. limit as weight increases.
(vii) In the measurement of stall speeds, the lowest test altitude is usually dictated by flig ht test safety concerns. This test data must then be expanded to lower altitudes, and hence lower Mach numbers, to cover the operational envelope of the airplane. Since C LMAX increases as the Mach number is reduced, simple expansion of the flight test data could result in extrapolating to a higher C LMAX than tested. Expansion of C LMAX versus Mach number data is only permitted up to the highest C LMAX within the range of W/ δ’s tested.
(viii) Stalling speeds are then calculated using the equation:
VS(e) =
295.37(W) (C S)L Where: W = a series of weights chosen as the independent variable - lbs.
CL = stall CL corresponding to the chosen weight (see Figure 29-3).
S = reference wing area - ft. 2
(ix) A plot of stall speed ve rsus weight, showing each flap/gear configuration, is then constructed (see Figure 29-4).