FRSR +50 KNOTS FRSR +50 KNOTS PULL (+) VIRIM
STICK BREAKOUT
FORCE FORCE
1.4 VS1 + FRSR PUSH (->
b. Procedures. None required for this section.
2. § 23.173 Static Longitudinal Stability.
a. Explanation.
(1) Demonstration Conditions. The general requirements of § 23.173 are determined from a demonstration of static stability under the conditions specified in § 23.175.
(2) Control Frictions. Section 23.173(b) effectively limits the amount of control friction that will be acceptable since excessive friction would have a masking effect on stability.
If autopilot or stability augmentation systems are of such a design that they tend to increase the friction level of the longitudinal control system, critical static longitudinal stability tests should be conducted with the system installed. Control cable tensions should be set to the maximum.
Note: A procedure that could be used is shown in figure 21 below:
Vpe,M,c
PULL X X t (-r) X .. , I STICK
V, FORCE I I
1.3V51 ~I I
v., 1• FRSR 11 1 PUSH t (·) VNE X= FRSR +(>40 KNOTS OR 15% Vrn,Ml 11/16/2011 AC 23-8C Figure 21 – Static Longitudinal Stability Data over Complete Speed Range
(3) Stable Slope. Section 23.173(c) is extremely genera l. It requires the test pilot's best judgment as to whether the stable slope of the stick force curve versus speed is sufficiently steep so that perceptibility is satisfactory for the safe operation of the airplane.
(4) Maximum Allowable Speed. Should be taken to mean V FE, VLE, VNE, and VFC/MFC, as appropriate.
b. Procedures. Refer to the guidance for § 23.175.
3. § 23.175 Demonstration of Stat ic Longitudinal Stability.
a. Explanation. Section 23.175 requires that, for cruise configuration, static longitudinal stability tests be conducted at representative cruising speeds at high and low altitude up to VNE or VFC/MFC, as appropriate, except that the trim speed need not exceed VH.
Section 23.173(a) states that static longitudinal stability must be shown at any speed that can be obtained; therefore, the longitudinal stability demonstration must cover the entire range from VS1 to VNE or VFC/MFC. Figure 21 shows typical coverage of the speed range in cruise with overlapping data. Mid-range trim points should include speed for best endurance, range, and high-speed cruise.
(1) VS1. Trim at VS1 + (>40 kts or 15 percent) + an estimate of the free return speed range (FRSR), perform static longitudinal stability tests from the trim speed within the speed range ensuring that the aircraft does not stall.
(2) VH. Determine V H at lowest altitude at MCP, perform longitudinal static stability tests within the prescribed speed range but do not exceed VNE.
(3) Additional trim points. Select additional trim points, for example, speed for best range and endurance, and so forth, until the speed range covered by data, refer to figure 22.
(4) Highest operating altitude. Go to highest operating altitude, depending on pressurization, oxygen requirements, and so forth, trim at VH, and repeat the test to a maximum speed of VFC/MFC or VNE, whichever comes first. Note that a stable slope above VNE, or VFC/MFC, or outside the (>40 kts or 15 percent) is desired, but not required.
b. Procedures.
(1) Section 23.l75(a) Climb .
(i) Stabilized Method. The airplane should be tr immed in smooth air for the conditions required by the regulation. Tests should be conducted at the critical combinations of weight and c.g. Normally, light weight and aft c.g. are critical.
(a) After observing trim speed, apply a pull force and stabilize at a slower speed. Continue this process in increments of 10 to 20 knots, depending on the speed spread being investigated, until reaching minimum speed for steady unstalled flight. At some stabilized point, the pull force should be gradually relaxed to allow the airplane to slowly return toward 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 be somewhat less than the original trim speed. As required by § 23.173, the new speed, cal led free-return speed, must be within 10 percent (7.5 percent for commuter category airplanes in cruise) of the trim speed.
(b) Starting again at the trim speed, push forces should be applied and gradually relaxed in the same manner as previously described at speeds up to 115 percent of the trim speed, and the same determination should be made.
(c) The flight test data band should be ± 2,000 feet from the trim altitude to minimize changes in power/thrust with altitude at a fixed throttle setting that could affect static longitudinal stability. High performance airplane s in the climb configuration sometimes require a number of iterations to stay within the data band. Or if the airplane has a very high thrust to weight ratio, it may need 3,000 feet as used for part 25 airplanes (refer to AC 25-7).
(2) Section 23.175(b), Cruise
(i) Acceleration-Deceleration Method. The stabilized flight test technique described in paragraph (1)(i) above is suitable for low performance airplanes or airplane configurations with low climb performance. The acceleration-deceleration method is particularly suitable for airplanes with high cruise speed. The airplane is trimmed at the desired airspeed and the power/thrust setting noted. Powe r/thrust is then increased to accelerate the airplane to the extreme speed of the desired data band. The power/thrust is then reset to the original trim power setting and the airplane allowed to decelerate at a constant altitude back to the original trim speed. Longitudinal static stability data is obtained during the deceleration to trim speed with the power and the elevator trim position the same as the original trim data point.
The data below trim speed is obtained in a similar manner by reducing power to decelerate the airplane to the lowest speed in the data band, reset the power to trim conditions, and record the data during the level acceleration back to trim speed. If, because of thrust/drag relationships, the airplane has difficulty returning to the trim data point, small altitude changes within ± 2,000 feet can also be used to coax an airplane acceleration-deceleration back to trim speed, but level flight 84
is preferred if possible. The data to be measured approximately every ten knots are speed and elevator stick force.
(3) Other Stability Test Procedures. The balance of the static longitudinal stability requirements are flown using either the stabilized method or the acceleration-deceleration method, but using the configurations, trim points, and speed ranges prescribed in § 23.175.
c. Data Acquisition and Reduction. Force readings can be made with a hand-held force gauge, fish scale, or by electronic means, and plotted against calibrated airspeed to determine compliance with the regulation. Refer to figure 22 for an example of the data plot.
Collect test data within a reasonable altitude band of the trim point altitude, such as ±2,000 feet.
Stick force measurements must be made unless:
(1) Changes in speed. Changes in speed are clearly reflected by changes in stick forces; and
(2) Maximum forces. The maximum forces obtained under §§ 23.173 and 23.175 are not excessive.
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11/16/2011 AC 23-8C Figure 22 - Static Longitudinal Stability Plot (Cruise Condition) 20 15STICK FORCE (POUNDS) 10 PULL 5 0 5 PUSH 10 70 90 110 130 150 170 190
1.3 VS1 VNE
CALIBRATED AIRSPEED (KNOTS)
4. § 23.177 Static Direction and Lateral Stability.
a. Explanation.
(1) Purpose. The purpose of this section is to require positive dire ctional and lateral stability and to verify the absence of rudder lock tendencies.
(2) Directional Stability. In § 23.177(a), the determination of "appropriate" wings level sideslip (previously referred to as skid) angles will depend on sound judgment in considering such things as airplane size, maneuverability, control harmony, and forces to determine the magnitude of wings level sideslip angles that the airplane will probably experience in service. Tests are conti nued beyond these "appropriate" angles up to the point where full rudder control is used or a force limit of 150 pounds, as specified in § 23.143, is reached. The rudder force may lighten, but may not reverse. The rudder force tests are conducted at speeds between 1.2 VS1 and VA. The directional stab ility tests are conducted at speeds from 1.2 VS1 to VNE or the maximum allowable speed for the configuration, whichever is limiting.
(3) Lateral Stability (Dihedral Effect). The static lateral stability tests (reference § 23.177(b)) take a similar approach since the basic requirement must be met up to maximum sideslip angles "appropriate to the type of airplane." Up to this angle, the airplane must demonstrate a tendency to raise the low wing when the ailerons are freed. The static lateral stability may not be negative at 1.2 VS1 in the takeoff configuration and 1.3 VS1 in other configurations.
(4) Forces. The requirement of § 23.177(d) is to be tested at a speed of 1.2 VS1 and larger than "appropriate" sideslip angles. At angles up to those that require full rudder or aileron 86
control, or until the rudder or aileron force limits specified in the table in § 23.143 are reached, the aileron and rudder force may lighten but may not reverse.
(5) Maximum Allowable Speed. Should be taken to mean V FE, VLE, VNE, and VFC/MFC, as appropriate.
(6) Autopilot or Stability Augmentation Systems (SAS). If autopilot or SAS are of such a design that they tend to increase the friction levels of the lateral and directional controls systems, then critical lateral and directional tests should be conducted with those systems installed, but not operating.
b. Procedures.
(1) Altitude. The tests should be conducted at the highest practical altitude considering engine power and aerodynamic damping.
(2) Loading. The airplane should be loaded to a nominal condition. Large lateral imbalance may mask lateral stability.
(3) Directional. To check static directional stabi lity with the airplane in the desired configuration and stabilized on the trim speed, the airplane is slowly yawed in both directions keeping the wings level with ailerons. When the rudder is released, the airplane should tend to return to straight flight. Some deadband around neutral is acceptable. Refer to paragraph
23.161. for guidance on ground adjustable tabs.
(4) Lateral. To check lateral stabil ity with a particular configuration and trim speed, conduct sideslips at the trim speed by maintaining the airplane's heading with rudder and banking with ailerons. Refer to paragraph 23.161 for guidance on ground adjustable tabs.
Section 23.177(b) requires the slip angle to be appropriate to the type of airplane and the bank angle to be at least l0 degrees. Some airplane s cannot maintain a heading in a slip with a l0 degree bank angle. In those cases , the slip should be performed with no less than a l0 degree bank and full opposite rudder and the heading allowed to vary. When the ailerons are released, the low wing should return to level. Some d eadband around neutral is acceptable. The pilot should not assist the ailerons during this evaluation. The pilot should hold the rudder as the wings level. For cases where there is not enough rudder to hold a steady heading, with bank angle, the pilot should use full rudder until the bank angle lowers to that where a steady heading can be maintained, then the rudder should be modulated as required to maintain the steady heading as the wings level.
c. Data Acquisition. Data recorded should be sufficient for showing compliance.
5. § 23.179 Stability. Reserved.
6. § 23.181 Dynamic Stability.
a. Explanation. Longitudinal Dynamic Stability.
(1) Short and Long Period Modes. Most normally configured airplanes will exhibit two distinct longitudinal modes of motion. The short period mode is the first response experienced after disturbing the airplane from its trim condition with the elevator control. It involves a succession of pitch acceleration, pitch rate, and pitch attitude changes that occur so rapidly the airspeed does not change significantly. Angle-of-attack will change in response to the pitching motions and produce accompanying changes in normal acceleration. Vertical gusts 87
and configuration changes, such as deploying flaps or speed brakes, may also excite the short period mode. The influence of control system springs and bob weights can be significant. If the disturbance from the trim condition is sustained long enough for the airspeed to change significantly, and if the pitch attitude excursions are not constrained by the pilot, the long period (or phugoid) oscillation will be excited with large but relatively slow changes in pitch attitude, airspeed, and altitude.
(2) Damping. Both the short period and long period modes are normally oscillatory in nature. However, the short period motion tend s to be so heavily damped that no significant overshoot or residual oscillations are perceptible to the pilot, a condition described qualitatively as "deadbeat.” If this is not the case, it shoul d be determined that the motions do not interfere with performance of any required maneuver or task. The long period, or phugoid oscillation, is characteristically lightly damped, sometimes even unstable. Mild levels of instability are acceptable as long as they do not significantly interfere with normal piloting tasks such as trimming to a desired speed, holding altitude, or glide slope tracking. Useful guidelines include:
the oscillation should be near neutrally stable if the period is less than 15 seconds or, for motions with longer period, the time to double amplitude should be greater than 55 seconds.
b. Procedures. Longitudinal, Short Period.
(1) General. The test for short period longit udinal dynamic stability is accomplished by a movement of pulse of the longitudinal control at a rate and degree to obtain a short period pitch response from the airplane. Initial inputs should be small and conservatively slow until more is learned about the airplane’s response. Gradually, the inputs can be made large enough to evaluate more readily the airplane’s oscillatory response and number of overshoots of the steady state condition.
(2) The Doublet Input. The “doublet input” excite s the short period motion while suppressing the phugoid. It is generally consider ed to be the optimum means of exciting the short period motion of any airplane. The doublet input causes a deviation in pitch attitude in one direction (nose down) then cancels it with a deviation in the other direction (nose up). The total deviation in pitch attitude from trim at the end of a doublet is zero. Thus, the phugoid mode is suppressed. However, the short period motion will be evident since the doublet generates deviation in pitch rate, normal acceleration, and angle-of-attack at a constant airspeed. Short period characteristics may be determined from the manner in which these parameters return to the original trimmed conditions. The doublet is performed as follows:
(i) Flight Condition. Stabilize and trim carefully in the desired configuration at the desired flight condition.
(ii) Control Inputs. With a smooth but fairly rapid motion, apply airplane nosedown longitudinal control to decrease pitch attitude a few degrees then reverse the input to noseup longitudinal control to bring the pitch attitude back to trim. As pitch attitude reaches trim, return the longitudinal cockpit control to trim and release it (controls-free short period) or restrain it in the trim position (controls-fixed short period). Both methods should be utilized. At the end of the doublet input, pitch attitude should be at the trim position (or oscillating about the trim position) and airspeed should be approximately trim airspeed.
(iii) Short Period Data. Obtaining quantitative information on short period characteristics from cockpit instruments is difficult and will be almost impossible if the motion is heavily damped. Short period oscillations are of ten of very low amplitude. If the pilot cannot see enough of the motion to measure and time a half-cycle amplitude ratio, the short period motion should be qualitatively described as essentially deadbeat.
(iv) Input Frequency. The frequency with which the doublet input is applied depends on the frequency and response characteristics of the airplane. The test pilot should adjust the doublet input to the particular airplane. The maximum response amplitude will be generated when the time interval for the complete doublet input is approximately the same as the period of the undamped short period oscillation.
(v) Sequence of Control Inputs. The doublet input may be made by first applying aft stick then reversing to forward stick. However, this results in less than 1g normal acceleration at the completion of the doublet and is more uncomfortable for the pilot.
(3) The Pulse Input. The pulse input also excites the short period nicely; however, it also tends to excite the phugoid mode. This confuses data analysis since the response of the airplane through the phugoid may be taken as a part of the short period response. This is particularly true for low frequency, slow-responding airplanes. Therefore, the pulse can usually only be utilized for high frequency, quick-responding airplanes in which the short period motion subsides before the phugoid response can develop. The pulse can always be used for a quick, qualitative look at the form of the short period motion. It is performed as follows:
(i) Flight Condition. Stabilize and trim in the desi red configuration at the desired flight condition.
(ii) Control Inputs. With a smooth but fairly rapid motion, apply airplane noseup longitudinal control to generate pitch rate, normal acceleration, and angle-of-attack changes then return the longitudinal control stick to the trim position. The short period motion may then be observed while restraining the control stick at the trim position (controls-fixed short period) or with the control stick free (controls-free short period).
(iii) Sequence of Control Inputs. Pulses may also be performed by first applying airplane nose-down longitudinal control.
(4) Conditions and Configurations. Short period dynamic longitudinal stability should be checked under all the conditions and configurations where static longitudinal stability is checked. Therefore, the test p ilot may find it convenient to test for both on the same flights. It is not intended nor required that every point along a stick force curve be checked for dynamic stability; however, a sufficient number of points should be checked in each configuration to ensure compliance at all operational speeds.
c. Procedures. Longitudinal Long Period (phugoid ) Dynamic Stability.
(1) General. The test for the phugoid mode is accomplished as follows:
(i) Cause the airplane to depart a significant amount from trim speed (about +10 percent should be sufficient) with an elevator input; and
(ii) Allow the ensuing oscillations in speed, rate of climb and descent, altitude, and pitch attitude to proceed without attempting to constrain any of the variables as long as airspeed, load factor, or other limitations are not exceeded.
(2) The Pulse Input. An appropriate control input for the phugoid test is a relatively slow elevator pulse to cause the airplane to increase or decrease speed from the trim point. Once the speed deviation is attained, the control is moved back to the original position and released.
(3) Conditions and Configurations. Long period dynamic stability should be checked under all of the conditions and configurations for which longitudinal static stability is 89
checked. As in the short period case, it is not intended that every point along a stick force curve be checked for phugoid damping; however, enough conditions should be checked to determine acceptable characteristics at all operational speeds.
(4) Data. The phugoid motion proceeds slowly e nough that it is reasonable to record minimum and maximum airspeed excursions as a function of time and enable construction of an envelope from which time to half-double amplitude may be determined.
d. Explanation. Lateral/Directional Dynamic Stability. Characteristic lateraldirectional motions normally involve three modes:
(1) Roll mode. A highly-damped convergence, called the roll mode, through which the pilot controls roll rate and, hence, bank angle;
(2) Spiral. A slow-acting mode, called the spiral, that may be stable but is often neutrally stable or even mildly divergent in roll and yaw; and
(3) Dutch roll. An oscillatory mode, called the Dutch roll, that involves combined rolling and yawing motions and that may be excited by either rudder or aileron inputs or by gust encounters.
(a) In addition, short period yawing oscillations due to rudder floating may sometimes be observed. The roll mode will almost always be satisfactory as judged by the ability to precisely control bank angle and counter gust upsets unless the response is slowed by high roll inertia or inadequate roll control power. Section 23.181(b) requires that the Dutch roll mode be investigated and determined to damp to 1/10 amplitude within seven cycles below 18,000 feet and 13 cycles from 18,000 feet to the certified maximum altitude. Also, any short period yawing oscillation associated with rudder motions must be heavily damped.
e. Procedures. Lateral/Directional. Two of the methods that may be used are described below:
(1) Rudder Pulsing. The rudder pulsing technique excites the Dutch roll motion nicely while suppressing the spiral mode if performed correctly. In addition, this technique can be used to develop a large amplitude oscillation that aids in data gathering and analysis, particularly if the Dutch roll is heavily damped. It is performed as follows:
(i) Flight Condition. Stabilize and trim carefully in the desired configuration at the desired flight condition.
(ii) Control Inputs. Smoothly apply alternating le ft and right rudder inputs in order to excite and reinforce the Dutch roll motion. Restrain the lateral cockpit control at the trim condition or merely release it. Continue the cyclic rudder pulsing until the desired magnitude of oscillatory motion is attained, then smoothly return the rudder pedals to the trim position and release them (controls free) or restrain them (controls fixed) in the trim position.
(iii) Input Frequency. The frequency with which the cyclic rudder inputs are applied depends on the frequency and response characteristics of the airplane. The test pilot should adjust the frequency of rudder pulsing to the particular airplane. The maximum Dutch roll response will be generated when the rudder pulsing is in phase with the airplane motion and the frequency of the rudder pulses is approximately the same as the natural (undamped) frequency of the Dutch roll.
(iv) Spiral Motion. The test pilot should attempt to terminate the rudder pulsing so that the airplane oscillates about a wings-level condition. This should effectively suppress the spiral motion.
(v) Data. Obtaining quantitative informati on on Dutch roll characteristics from cockpit instruments and visual observations requires patience, particularly if the motion is heavily damped. If instrumentation is available to record sideslip angle versus time, the dynamic characteristics of the maneuver can readily be determined. The turn needle of the needle-ball instrument can also be used to observe 1/10 amplitude damping and the damping period.
(2) Steady Sideslip. The steady sideslip release can also be used to excite the Dutch roll; however, the difficulty in quickly returning the controls to trim and the influence of the spiral mode often precludes the gathering of good quantitative results. Full rudder or a large amplitude sideslip may cause high loads on the airplane. The rudder pulsing technique usually produces better Dutch roll data. The steady sideslip release technique is performed as follows:
(i) Flight Condition. Stabilize and trim carefully in the desired configuration at the desired flight condition.
(ii) Control Input. Establish a steady heading sides lip of a sufficient magnitude to obtain sufficient Dutch roll motion for analysis. Utilize maximum allowable sideslip, using rudder as required. Stabilize the sideslip carefully. Quickly, but smoothly, return all cockpit controls to trim and release them (controls-free Dutch roll) or restrain them at the trim position (controls-fixed Dutch roll). Bo th methods should be utilized.
f. Stability Augmentation Systems (SAS). If the airplane is equipped with SAS, the airplane's characteristics should be evaluated throughout the approved operating envelope following failures that affect the damping of the applicable mode. Following a SAS failure, if unsatisfactory damping is confined to an avoidable flight area or configuration and is controllable to return the airplane to a satisfactory operational condition for continued safe flight, the lack of appreciable positive damping may be acceptable. Control of the airplane, including recovery, should be satisfactory using applicable control inputs. Following a critical failure, the degree of damping required should depend on the effect the oscillation will have on pilot tasks, considering environmental conditions. The capability to handle this condition should be demonstrated and evaluated. If a satisfactory reduced operational envelope is developed, appropriate procedures, performance, and limitations should be placed in the AFM. If a critical failure results in an unsafe condition, a redundant SAS may be required. The FAA has granted relief from the Dutch roll criteria for high performance jets that operate at high altitudes. These exemptions essentially allow for reduced damping until the pilots can decend to an altitude where the jet can aerodynamically meet the damping requirements in part 23. Examples of the exemptions are available on the FAA’s website.
g. Data Acquisition and Reduction. Data acquisition for this test should support a conclusion that any short period oscillation is heavily damped and any Dutch roll is damped to 1/10 amplitude in seven cycles below 18,000 feet and 13 cycles from 18,000 feet to the certified maximum altitude.
h. Maximum Allowable Speed. Should be taken to mean V FE, VLE, VNE, and VFC/MFC, as appropriate.
Section 7. Stalls
1. § 23.201 Wings Level Stall.
a. Explanation.
(1) Stall. Section 23.201(b) defines when the ai rplane can be considered stalled for airplane certification purposes. When one of three conditions occurs, whichever occurs first, the airplane is stalled. The conditions are as follows:
(i) Uncontrollable downward pitching motion;
(ii) Downward pitching motion that results from the activation of a device (for example, stick pusher); or
(iii) The control reaches the stop.
Note: The term "uncontrollable downward pitching motion" is the point at which the pitching motion can no longer be arrested by application of nose-up elevator and not necessarily the first indication of nose-down pitch. Figure 1 shows a graphic representation of stall speed time histories for various configurations.
(2) Related Sections. The stalled condition is a flight condition that comes within the scope of §§ 23.49, 23.141, 23.143(b), 23.171, and 23.173(a). Section 23.143(b) requires that it be possible to effect a "smooth transition" from a flying condition up to the stalled flight condition and return without requiring an exceptional degree of skill, alertness, or strength. Any need for anticipated or rapid control inputs exceeding that associated with average piloting skill is considered unacceptable.
(3) Recovery. The flight tests include a determination that the airplane can be stalled, and flight control recovered, with normal use of the controls. Section 23.201(a) requires that it must be possible to produce and correct roll by unreversed use of the roll control and to produce and correct yaw by unreversed use of the directional control. The power used to regain level flight may not be applied until flying control is regained. This is considered to mean not before a speed of 1.2 VS1 is attained in the recovery dive.
(4) Power.
(i) Power off. The propeller condition for the "power-off" tests prescribed by § 23.201(e)(4) should be the same as the "throttles closed" condition prescribed for the stalling speed tests of § 23.49, that is, propellers in the takeoff position, engine idling with throttles closed. The alternative of using sufficient power to produce zero thrust does not apply to stall characteristics demonstrations.
(ii) Power on. For the power-on tests, an ex treme nose up attitude is normally considered to be a pitch attitude of more than 30 degrees.
(5) Configurations. Stall characteristics should be evaluated as follows:
(i) At maximum to minimum weights at aft c.g., aft light loadings may be the most critical in airplanes with high thrust-to-weight ratios;
(ii) With the elevator up stop set to the maximum allowable deflection;
(iii) With maximum allowable lateral imbalance; and
(iv) At or near maximum approved altitude for airplanes approved to fly up to and above 25,000 feet. Clean, and any other reasona ble configurations (such as spoiler), are the only stalls necessary at high altitude.
Note: Airplanes with de-rated engines s hould be evaluated up to the critical altitude of the engine and at maximum altitude for which the airplane is to be certified. An airplane may be approved if it has stick pusher operation in one configuration, such as power on, and has acceptable stall characteristics for the remaining configurations.
b. Procedures.
(1) Emergency Egress. It is the responsibility of the applicant to provide adequate provision for crew restraint, emergency egress, and use of parachutes (reference § 21.35(d)).
(2) Buildup. Generally, the stalls at more rearward, c.g. positions are more critical than at the forward c.g. position. For this reason, the stall characteristic s at forward c.g. should be investigated first. Altitude should be low enough to ensure capability of setting 75 percent power, but high enough to accomplish a safe recovery. The 75 percent power requirement means 75 percent of the rated power adjusted to the temperature and altitude test conditions.
Reciprocating engine tests conducted on a hot day, for example, would require higher manifold pressures to be set so that when chart brake horsepower is adjusted for temperature the result is 75 percent power.
(3) Pilot Determinations. During the entry and recovery, the test pilot should determine the following:
(i) That the stick force curve remains positive up to the stall (that is, a pull force is required) (reference § 23.171) when the trim speed is higher than the stall speed.
(ii) That it is possible to produce and correct roll and yaw by unreversed use of the rolling and directional control.
(iii) The amount of roll or yaw encountered during the recovery.
(4) Speed Reduction Rate. Section 23.201(b) requires th e rate of speed reduction for entry not exceed one knot per second.
c. Data Acquisition and Reduction.
(1) Instruments. The applicant should provide a recently calibrated sensitive altimeter, airspeed indicator, accelerometer, outside air temperature gauge, and appropriate propulsion instruments; such as a torque meter or manifold pressure gauge and tachometer, a means to depict roll, pitch, and yaw angles; and force gauges when necessary.
(2) Data Recording. Automatic data recording is desirable, but not required, for recording time histories of instrumented parameters and such events as stall warning, altitude loss, and stall break. The anal ysis should show the relationship of pitch, roll, and yaw with respect to various control surface deflections. (Refer to figure 1, stall speed determination.)