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.)
d. Stick Pusher.
(1) Background. Stick pushers have been installe d in some airplanes that would not meet the requirements of § 23.201. This was accomplished under the provisions of § 21.21b(1).
In some airplanes, operation of the stick pusher was not critical to safe flight and, in others, stick pusher performance was essential to safe flight. In the latter case, the stick pusher typically functions as a stall barrier to prevent an airplane from entering flight regimes where a nonrecoverable stall could occur.
(2) Stall Prevention. There are two basic situations where a stick pusher would be necessary to show compliance with regulations. These are the following:
(i) Airplane Recoverable. The stall characteristics are investigated and, during these tests, the airplane does not meet regulatory requirements but an inadvertent aerodynamic stall would not be catastrophic or require exceptional piloting skill to recover. For example, during certain conditions of c.g., weight, or power, the airplane exceeds the l5 degree roll limit because of excessive angle-of-attack or other conditions. The pusher is installed and designed to function at some angle-of-attack where the 15 degrees of roll will not be exceeded. If this pusher system fails to function, the airplane is still recoverable from the stalled flight condition. For this system, the occurrence of a failure should be evaluated for an unsafe condition.
(ii) Airplane Not Recoverable. When the airplane is not capable of recovering from stalls or the applicant chooses not to investigate the consequences of stalling demonstrations with the pusher system rendered inoperative, then if the stick pusher fails to perform its intended function, an unsafe condition would exist.
(3) System Tolerances. Stick pusher system(s) tolera nces should be evaluated and accounted for during certification flight tests. The applicant normally specifies the system tolerances in terms of plus or minus so many knots. The system(s) should be set to the minus (lowest stall speed) side of the tolerance when investigating stall characteristics and minimum longitudinal trim capability. The system(s) should be set to the plus (highest stall speed) side of the tolerance for stall speed determination and for determining performance stall speed multiples.
Alternately, a nominal stick pusher stall speed may be used when it is determined that stick pusher tolerances result in no more than three knots or five percent variation in stall speed, whichever is greater.
(4) Airspeed Margins. The airspeed margin between unsatisfactory stall characteristics and the minimum stick pusher actuation speed, for identical flight conditions, should be evaluated. The following information is provided as a guide. For airplanes with unsatisfactory, hazardous or unrecoverable aerodynamic stall characteristics, the minimum speed margin between aerodynamic stall and minimum stick pusher systems actuation speed should not be less than five knots. For other airplanes with known and less hazardous aerodynamic stall characteristics, the speed margin may be reduced to not less than two knots.
(5) Preflight Check. If a reliability credit is to be given for a preflight check, the following should be evaluated:
(i) The check includes the functioning of the complete system, including the incidence sensors, so all faults would be detected.
(ii) The check is easily conducted, and requires little pilot time or effort.
(iii) A note in the limitations section of the AFM requires the check to be accomplished prior to flight.
(iv) The AFM identifies the criticality of the system and the need to accomplish the preflight check.
(6) Inadvertent Operation. Inadvertent stick pusher operation should be extremely improbable or investigated and shown not to be hazardous and to be recoverable.
2. § 23.203 Turning Flight and Accelerated Turning Stalls.
a. Explanation.
(1) Wings level stalls. Explanations for wings level stal ls also apply to turning flight and accelerated turning stalls.
(2) Turning flight and accelerated turning stalls. The only differences between the investigation required for turning flight and accelerated turning stalls are in the speed reduction rate and the accepted roll off bank angles.
b. Procedures.
(1) Procedures for wings level stalls applies to turning flight and accelerated turning stalls.
(2) Test pilot determinations. During the maneuver, the test pilot should determine the following:
(i) That the stick force remains positive up to the stall;
(ii) That the altitude lost is not, in the test pilot's opinion, excessive;
(iii) There is no undue pitch up;
(iv) That there are no uncontrollable spinning tendencies; that is, while the airplane may have a tendency to spin, a spin entry is readily preventable;
(v) That the test pilot can complete the recovery with normal use of the controls and average piloting skill;
(vi) Roll does not exceed the value specified in the requirements; and
(vii) For accelerated turning stalls, maximum speed or limit load factors were not exceeded.
(3) Section 23.203(a). Section 23.203(a) requires the rate of speed reduction for a turning flight stall not exceed one knot per second; for an accelerated turning stall, three to five knots per second with steadily increasing normal acceleration.
c Data Acquisition. Same as for wings level stalls.
3. § 23.205. Reserved.
4. § 23.207 Stall Warning.
a. Explanation.
(1) Purpose. The purpose of this requirement is to ensure an effective warning in sufficient time to allow a pilot to recover from an approach to a stall without reaching the stall.
(2) Types of Warning. The effective warning may be from either aerodynamic disturbances or from a reliable artificial stall-warning device such as a horn or a stick shaker.
The aerodynamic warning is usually manifested by a buffet, which vibrates or shakes the airplane. The type of warning should be the same for all configurations.
(3) Artificial Stall Warning. Stall warning devices may be used in cases where there is inadequate aerodynamic warning. The warning signal from the devices should be clear and distinctive and not require the pilot's attention to be directed inside the airplane. A stall warning light by itself is not acceptable. If a stick shaker is installed, the warning should be unmistakable even if flying hands off.
(4) Margin. The stall warning margin, as defined in the rule, is applicable with speed reduced at a rate of one knot per second for §§ 23.201 and 23.203(a)(1).
b. Procedures. The stall warning tests should be conducted in conjunction with the stall tests required by §§ 23.201 and 23.203.
Section 8. Spinning
1. § 23.221 Spinning.
a. Explanation.
(1) Spin. A spin is a sustained autorotation at angles-of-attack above stall. The rotary motions of the spin may have oscillations in pitch, roll, and yaw superimposed upon them.
The fully developed spin is attained when the trajectory has become vertical and the spin characteristics are approximately repeatable from turn to turn. Some airplanes can autorotate for several turns, repeating the body motions at some interval, and never stabilize. Most airplanes will not attain a fully developed spin in one turn.
(2) Category Spins. Section 23.221 addresses three situations:
(i) Normal category spins.
(ii) Utility category spins.
(iii) Acrobatic category spins.
(3) Utility Category Airplanes. Utility category is used for airplanes intended for limited acrobatic operations in accordance with § 23.3. Spins (if approved for the particular type of airplane) are considered to be a limited acrobatic operation. This type of airplane may be approved in accordance with § 23.221(a), normal category, or with § 23.221(c), acrobatic category.
(4) Stick Pushers. Single engine airplanes equipped with a stick pusher will not be required to demonstrate compliance with § 23.221; however, the applicant will need to document the equivalency using an equivalent level of safety finding.
(5) Margin. The stall warni ng margin, as defined in the rule, is applicable when speed is reduced at a rate of one knot per second for §§ 23.201 and 23.203(a)(1).
b. Guidance and Procedures Applicable to Both Normal and Acrobatic Category Spins.
(1) Weight and C.G. Envelope. Refer to paragraph 23.23 of this AC for guidance on weight and c.g. envelope exploration.
(2) Moments of Inertia. Moments of inertia should also be considered when evaluating the c.g. envelope. Most general aviati on airplanes have low inertias combined with high aerodynamic damping and relatively similar moments of inertia along the wing and fuselage axis. However, designs of modifications, such as wingtip fuel tanks, can change the spin recovery time and possibly the recovery method. Applicants are encouraged to consider these effects and approach flight testing at extreme mass distributions with caution.
(3) Control Deflections. Control surface deflections, i. e., rigging, should be set to a nominal value based on the production specifications. If the test results suggest some peculiarities that would indicate compliance issues, then the critical side of the allowable tolerances should be tested. For example, a possible spin flight test program could be to perform the spin matrix with the controls set at the nominal deflection values. Analysis of the data may show the critical conditions for entry and recovery. If a critical condition is found, defined, and agreed upon by the FAA, these critical tests are repeated with the control deflections set to the most critical tolerances.
(4) 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)).
(5) Spin Recovery Parachutes.
(i) Spin recovery parachutes should be installed on all airplanes requiring spin testing for certification unless experience with the airplane shows, and the applicant and the FAA agree that it is not necessary.
(ii) The spin recovery system installation should be carefully evaluated to determine its structural integrity, reliability, susceptibility to inadvertent or unwanted deployment or jettison, and adequate or redundant jettison capability. NASA recommendations should be referred to when evaluating the design of the chute deployment and jettison systems.
The chute type, diameter, porosity, riser length, and lanyard length should be determined in accordance with NASA recommended practices to maximize the probability the chute will be effective in spin recovery. Chute sizes, and particularly riser and lanyard lengths, depend strongly on such aircraft variables as wing design, fuselage shape, tail arm, and mass properties.
The sizes and lengths shown in the referenced NASA reports are for particular aircraft that were tested in the NASA Langley Spin Tunnel and will not necessarily be the correct size to recover other aircraft, even if the aircraft layout is similar. The following NASA documents are available from the National Technical Information Service (NTIS); 5285 Port Royal Road;
Springfield, Virginia 22161; or at the following email address: http://www.ntis.gov/ Appropriate NASA recommendations can be found in the following publications:
(a) NASA Technical Paper 1076, "Spin-Tunnel Investigation of the Spinning Characteristics of Typical Single-Engine General Aviation Airplane Designs," dated November 1977.
(b) NASA Technical Note D-6866, "Summary of Design Considerations for Airplane Spin-Recovery Parachute Systems."
(c) NASA Conference Paper, CP-2127, l4th Aerospace Mechanisms Symposium, May 1980, entitled, "A Spin-Recovery System for Light General Aviation Airplanes."
(iii) Final certification of the spin characteristics should be conducted with the external spin chute removed unless it is determined that spin chute installation has no significant effect or a conservative effect on spin characteristics.
(6) Build-Up. When any doubt exists regarding the recovery characteristics of the test airplane, a build-up technique should be employed consisting of spin entries and recoveries at various stages as the maneuver develops. Excessive aerodynamic control wheel back pressure indicates a possibility of unsatisfactory spin characteristics. Any control force lightening or reversal is an indication of possible deep stall entry. Refer to paragraph c(7) for definition of excessive back pressure. A yaw ra te instrument is valuable in detecting progress toward a fully developed spin condition or an uncontrollable maneuver. Unusual application of power or controls has sometimes been found to induce uncontrollable spins. Leading with elevator in recovery and cutting power as the airplane rolls into a spin have been known to induce uncontrollable spins.
(7) Entry. Spins should be entered in the same manner as the stalls in §§ 23.201 and
23.203 with trim at 1.5 VS1 or as close as practical. As the ai rplane stalls, with ailerons neutral, apply full-up elevator and full rudder in the direction of spin desired.
(8) Recovery. Recoveries should consist of th rottle reduced to idle, ailerons neutralized, full opposite rudder, followed by forward elevator control as required to get the wing out of stall and recover to level flight. Fo r acrobatic category spins, the manufacturer may establish additional recovery procedures, provided they show compliance for those procedures with this section.
(9) Altitude. The effect of altitude should be investigated.
(10) Initial Investigation. In all cases, the initial spin investigation should be accomplished at as high an altitude above the ground as reasonably possible and a predetermined, pre-briefed "hard" altitude established to be used as the emergency egress altitude. In other words, if th e airplane cannot be recovered by that altitude, all occupants should exit the airplane without hesitation. The altitude selected should take into account the opening characteristics of the parachutes, the difficulty of egress, the estimated number of turns to get out and the altitude loss per turn, the distance required to clear the airplane before deploying the parachutes, and so forth.
(11) Power. The use of power during spin en try for both normal and abnormal spins is recommended in order to determine the effects of power on spin entry characteristics and possible spin recovery procedures. The power on setting should be that used for the power on demonstrations in paragraphs 23.201 and 23.203. For power-on normal category spins, the throttle can be reduced to idle after one turn.