b. Procedures. The procedures applicable to § 23.75 would apply for § 23.153 except that, for turbopropeller airplanes, the flight-idle fuel flow should be adjusted to provide minimum thrust.
51. SECTION 23.155 ELEVATOR CONTROL FORCE IN MANEUVERS.
a. Explanation.
(1) Stick Force Per G. The purpose of this requirement is to ensure that the positive stick force per g levels in a cruise configuration are of sufficient magnitude to prevent the pilot from inadvertently over-stressing the airplane during maneuvering flight. The minimum maneuvering stability levels are generally found at aft c.g. loadings. Both aft heavy and aft light loadings should be considered. During initial inflight investigations, caution should be exercised in the event that pitch-up tendencies or decreasing stick force per g conditions occur.
(2) Buffet Boundaries. Low speed buffet onset may occur during high altitude investigations. A qualitative evaluation should be conducted beyond the boundary of buffet onset to ensure a capability to maneuver out of the buffet regime.
b. Procedures. Compliance with the requirements of § 23.155 may be demonstrated by measuring the normal acceleration and associated elevator stick force in a turn while maintaining the initial level flight trim speed. A descent may be required in the turn to maintain the level flight trim speed. As a minimum, the following conditions should be investigated in the cruise configuration, that is, flaps up and gear up (if retractable):
Condition POWER Wings Level Trim Speed Altitude 1 See Note Trimmed (but not to exceed VNE or
VMO/MMO) Low 2 See Note Trimmed Altitude for highest dynamic pressure (q) 3 See Note VA Low 4 See Note VA Highest attainable approved altitude
Note: 75 percent maximum continuous power or maximum power (reciprocating engine) or maximum cruise power (turbine).
TABLE 51-1 - Procedures
Compliance may be demonstrated by measuring the normal acceleration achieved with the limiting stick force (50 lbs. for wheel controls, 35 lbs. for stick controls) or by establishing the stick force per g gradient and extrapolating to the appropriate limit. Linear stick force gradients may be extrapolated up to 0.5g maximum. Non-linear stick force gradients that indicate a possible gradient lightening at higher g levels should not be extrapolated more than 0.2g.
c. Data Acquisition and Reduction. The following should be recorded for each test condition:
(1) Wt./c.g.;
(2) Pressure altitude;
(3) Outside air temperature;
(4) Engine power parameters;
(5) Trim setting;
(6) Elevator force;
(7) Normal acceleration at c.g.; and
(8) Gear/flap position.
The test data should be presented in stick force versus g plots. Figure 51-1 shows a sample plot.
Test results should be compared to the requirements of § 23.155(a).
0.51 .0 1.52 .0 2.53 .0 3.5 10 20 30 40 50 0 NORMAL ACCELERATION - G's LBS PULL
4.0 CAS= ________
c.g.= ________
FIGURE 51-1 - STICK FORCE PER G
d. Stick Force Per G, § 23.155(c). An increase in pull force should be required to produce an increase in normal acceleration throughout the range of required load factor and speed. Any reduction in control force gradient with change in load factor should not be so large or abrupt as to significantly impair the ability of the pilot to maintain control of normal acceleration and pitch rate. The local value of control force gradient should not be less than 3 lbs/g for stick-controlled 69
airplanes or 4 lbs/g for wheel-controlled airplanes. The elevator control force should increase progressively with increasing load factor.
Flight tests to satisfy the above must be performed at sufficient points to establish compliance with § 23.155(c) throughout the normal flight envelope. During these tests, the load factor should be increased until either:
(1) The intensity of buffet provides a strong and effective deterrent to further increase of load factor; or
(2) Further increase of load factor requires an elevator control force in excess of 150 lbs. for a wheel control or 125 lbs. for a stick control or is impossible because of the limitations of the control system; or
(3) The positive limit maneuvering load factor is achieved.
52. SECTION 23.157 RATE OF ROLL.
a. Explanation. The purpose of this requirement is to ensure an adequately responsive airplane in the takeoff and approach configuration.
b. Procedures.
(1) Bank Angle. The airplane should be placed in a 30-degree bank and rolled through an angle of 60 degrees. For example, with the airplane in a steady 30-degree left bank, roll through a 30-degree right bank and measure the time. Sections 23.157(b) and (d) should be accomplished by rolling the airplane in both directions.
(2) Controls. Sections 23.157(a) and (c) permit using a favorable combination of controls.
The rudder may be used as necessary to achieve a coordinated maneuver.
(3) Weight. The "W" in the formulas is the maximum takeoff weight.
53.-62. RESERVED.
Section 5. TRIM
63. SECTION 23.161 TRIM.
a. Explanation. The trim requirements ensure that the airplane will not require exceptional skill, strength, or alertness on the pilot's part to maintain a steady flight condition. The tests require the airplane to be trimmed for hands-off flight for the conditions specified. It should be noted that for single-engine airplanes, lateral-directional trim is required at only one speed, thus ground adjustable tabs are acceptable. For lateral-directional testing, the tabs may be adjusted for the test trim airspeed and readjusted for subsequent tests. For multiengine airplanes, directional trim is required for a range of speeds. If appropriate, lateral baggage loading and fuel asymmetry should be considered in this evaluation.
b. Procedures.
(1) Actuator Settings. Trim actuator travel limits should be set to the minimum allowable.
(2) Altitude and Power. Tests for trim should be conducted in smooth air. Those tests requiring use of maximum continuous power should be conducted at as low an altitude as practical to ensure attaining the required power.
(3) Weight and C.G. Longitudinal trim tests should be conducted at the most critical combinations of weight and c.g. Forward c.g. is usually critical at slow speeds and aft c.g.
critical at high speeds.
64.-69. RESERVED.
Section 6. STABILITY
70. SECTION 23.171 GENERAL.
a. Explanation.
(1) Required Stability. The stability portion of Part 23 is primarily concerned with static stability. No quantitative values are specified for the degree of stability required. This allows simple test methods or qualitative determinations unless marginal conditions are found to exist. The regulations merely require that the airplane be stable and that it have sufficient change in control force, as it is displaced from the trimmed condition, to produce suitable control feel for safe operation.
(2) Forces. The magnitude of the measured forces should increase with departure from the trim speed up to the speed limits specified in § 23.175 or up to the 40 lb. force limit specified in § 23.173. The stick force variation with speed changes should be stable. In other words, a pull force required to fly slower than trim and a push force required to fly faster than trim and the gradient should be clearly perceptible to the pilot at any speed between 1.3 VS1 and VNE or VFC/MFC. Figure 70-1 below shows an example of cruise configuration:
Vi STICK FORCE PUSH (-) PULL (+) FRSR FRSR +> 40 KNOTS OR 15%VTRIM FRSR +> 40 KNOTS OR 15%VTRIM
FIGURE 70-1 - Static Longitudinal Stability Data
Speed Range = Greater of + 40 knots (kts) or 15% VTRIM + free return speed range (FRSR)
At speeds below 1.3 VS1 for normal, utility, and acrobatic airplanes, and at speeds below
1.4 VS1 for commuter airplanes, the slope need not be stable, see Figures 70-2 and 70-3.
The pull forces can decrease in magnitude with speed decrease down to, but not including, the stall speed VS1; however, the pull force shall in no case fall below zero before the stall is reached.
Instrumented force measurements are required if there is any uncertainty in the qualitative assessment of the force gradients.
Vi STICK FORCE PUSH (-) PULL (+) FRSR +> 40 KNOTS OR 15% VTRIM VS1 1.3 VS1 VTRIM
FIGURE 70-2 - Low Speed Instabilities; Normal, Utility, and Acrobatic Airplanes
Vi STICK FORCE PUSH (-) PULL (+) FRSR +50 KNOTS VS1 1.4 VS1 VTRIM
FIGURE 70-3 - Low Speed Instabilities; Commuter Airplanes
b. Procedures. None required for this section
71. SECTION 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.
Vi STICK FORCE PUSH (-) PULL (+) X X VS1
1.3 VS1 XX
FRSR VTRIM VNE X= FRSR +(>40 KNOTS OR 15% VTRIM)
VFC/MFC
FIGURE 71-1 – Static Longitudinal Stability Data over Complete Speed Range
A procedure that could be used is as follows:
(3) Stable Slope. Section 23.173(c) is extremely general. 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 VFE, VLE, VNE, and VFC/MFC, as appropriate.
b. Procedures. Refer to paragraph 72.
72. SECTION 23.175 DEMONSTRATION OF STATIC 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 74
longitudinal stability demonstration must cover the entire range from VS1 to VNE or VFC/MFC.
Figure 72-1 shows typical coverage of the speed range in cruise with overlapping data. Midrange trim points should include speed for best endurance, range, and high-speed cruise.
(1) Trim at VS1 + (>40 kts or 15%) + 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) Determine VH at lowest altitude at maximum continuous power (MCP), perform longitudinal static stability tests within the prescribed speed range but do not exceed VNE.
(3) Select additional trim points, for example, speed for best range and endurance, and so forth, until the speed range covered by data, see figure 72-1.
(4) 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 is not required.
b. Procedures.
(1) Section 23.l75(a) Climb .
(a) Stabilized Method. The airplane should be trimmed 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.
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, called free-return speed, must be within 10 percent (7.5 percent for commuter category airplanes in cruise) of the trim speed.
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.
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 airplanes 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).
(b) Acceleration-Deceleration Method. The stabilized flight test technique described in Section (a) 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. Power/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 75
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 is preferred if possible. The data to be measured approximately every 10 knots would be speed and elevator stick force.
(2) 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. See figure 72-1 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 are clearly reflected by changes in stick forces; and
(2) The maximum forces obtained under §§ 23.173 and 23.175 are not excessive.
70 90 110 130 150 170 190 10 5 0 5 10 15 20
1.3 VS1 VNE
CALIBRATED AIRSPEED (KNOTS) PULL PUSH STICK FORCE (POUNDS)
FIGURE 72-1 - STATIC LONGITUDINAL STABILITY PLOT (CRUISE CONDITION)
73. SECTION 23.177 STATIC DIRECTIONAL AND LATERAL STABILITY.
a. Explanation.
(1) Purpose. The purpose of this section is to require positive directional 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 continued 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 stability 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 at the 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 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 VFE, 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 maximum allowable lateral imbalance should be maintained. Both low fuel and full fuel loadings should be evaluated for possible effects of fuel movement.
(3) Directional. To check static directional stability 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. See paragraph 63.a. for discussion of ground adjustable tabs.
(4) Lateral. To check lateral stability 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. See paragraph 63.a. for discussion of ground adjustable tabs. Section 23.177(b) 77
requires the slip angle to be appropriate to the type of airplane and the bank angle to be at least l0 degrees. Some airplanes 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 tend to return to level. The pilot should not assist the ailerons during this evaluation. The pilot should hold full rudder during the evaluation (either up to the deflection limit or to the force limit, whichever occurs first).
c. Data Acquisition. Data recorded should be sufficient for showing compliance.
74. SECTION 23.179 RESERVED.
75. SECTION 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 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 tends 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 should 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 longitudinal 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” excites the short period motion while suppressing the phugoid. It is generally considered 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 nose-down longitudinal control to decrease pitch attitude a few degrees then reverse the input to nose-up 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 often 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 desired configuration at the desired flight condition.
(ii) Control Inputs. With a smooth but fairly rapid motion, apply airplane nose-up 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 79