(2) Analysis of System. An analysis of the control system should be completed before conducting the loss of the primary lateral control test. On some airplanes, the required single lateral control system failure could result in loss of a rudder aileron interconnect and, perhaps, loss of directional control as well as the primary lateral control. The most critical linkage failure of the primary lateral control system must be considered.
(3) Procedures. The wording of the regulation sufficiently describes the maneuvers required to show compliance. The selection of altitudes, weights, c.g. position, lateral imbalance, and aircraft configurations to be flight tested by the FAA will depend on the study of the applicant’s flight test report and whether the aircraft has a lateral trim system. Use of the lateral trim system to maneuver the aircraft and to hold wings level during an actual or simulated landing flare is authorized to comply with § 23.147(c). Those aircraft that do not have a separate and independent lateral trim system could use the rudder or differential power of a twin engine aircraft to generate a sideslip that would produce a rolling movement to control the bank angle.
The use of rudder or asymmetric power to control bank angle implies that the aircraft exhibits lateral stability or dihedral effect. For those aircraft that use a rudder aileron interconnect to obtain lateral stability for which it is possible for a single failure in the primary lateral control system to disconnect the aileron rudder interconnect, compliance with § 23.147(c) must be performed without the interconnect. If compliance with § 23.147(c) can only be demonstrated with either flap, speed, power or procedure restrictions, or all of these, these restrictions should be noted in the AFM, in the emergency section.
(i) Altitude. An intermediate altitude and, if the airplane is approved for flight above 25,000 feet, an altitude near the maximum capability of the airplane needs to be tested.
The high altitude test is to determine controllability with decreased Dutch roll damping.
(ii) Weight. Maximum gross weight for a ll tests, except where otherwise described in subparagraph (iii) below.
(iii) C.G. For conventional configuration, § 23.1 47(a) the most aft c.g. is critical if the rudder is used to roll the airplane. For unconventional configurations, the most critical c.g.
must be used.
(iv) Lateral Imbalance. The maximum lateral imbalance for which certification is requested must be used when flight testing for compliance with § 23.147(c).
(v) Configuration, Power and Speed. Lateral controllability must be demonstrated with all practicable configurations and speeds. The maximum flaps used to demonstrate an actual or simulated landing need not be the maximum deflection possible.
(4) Data Acquisition. No special instrumentation is required.
4. § 23.149 Minimum Control Speed.
a. Background. Section 23.149 requires the minimum control speed to be determined.
Section 23.1545(b)(6) requires the airspeed indicator to be marked with a red radial line showing the maximum value of one-engine-inoperative minimum control speed. Section 23.1583(a)(2) requires that VMC be furnished as an airspeed limitation in the AFM. These apply only to multiengine airplanes. A different VMC airspeed will normally result from each approved takeoff flap setting. There are variable factors affecting the minimum control speed. Because of this, VMC should represent the highest minimum airspeed normally expected in service. The variable factors affecting VMC testing include:
(1) Engine Power. V MC will increase as power is increased on the operating engine(s). Engine power characteristics should be known and engine power tolerances should be accounted for.
(2) Propeller of the Inoperative Engine. Windmilling propellers result in a higher VMC than if the propeller is feathered. V MC is normally measured with propeller windmilling unless the propeller is automatically feathered or otherwise driven to a minimum drag position (for example, NTS-System) without requiring pilot action.
(3) Control Position. The value of V MC is directly related to the control surface travel available. Normally, V MC is based on available rudder travel but may, for some airplanes, be based on lateral control travel. A stall speed faster than VMC defines the limit in lieu of a traditional control deflection. For these reasons, VMC tests should be conducted with rudder and lateral (if applicable) controls set at minimum travel. In addition, rudder and lateral control deflections should be adjusted to the minimum production tolerances. If, during VMC tests, control force limits would be exceeded at full deflection, then a lesser deflection should be used so as not to exceed § 23.143 force limits.
(4) Weight and C.G. For rudder limited airplanes w ith constant aft c.g. limits, the critical loading for VMC testing is most aft c.g. and minimum weight. Aft c.g. provides the shortest moment arm relative to the rudder thus the least restoring moments with regard to maintaining directional control. VMC should be determined at the most adverse weight.
Minimum practical test weight is usually the most critical because the beneficial effect of banking into the operating engine is minimized. Light weight is also desirable for VMC testing because the stall speed is reduced.
(5) Lateral Loading. The maximum allowable adverse lateral imbalance (fuel, baggage, and so forth) should be maintained.
b. Explanation.
(1) Controllability. The determination of V MC is closely related to the controllability requirements. It is one of the maneuvers that generally requires either maximum rudder or near maximum aileron deflection (unless limited by temporary control forces), or both, to maintain airplane control. When minimum control speed is determined using maximum rudder deflection, limited airplane maneuvering is still available using the ailerons and elevator.
When minimum control speed is determined using near maximum aileron deflection, the airplane may be incapable of further maneuvering in the normal sense.
(2) Critical Engine. The regulation requires that V MC determination be made "when the critical engine is suddenly made inoperative." The intent is to require an investigation to determine which engine is critical from the standpoint of producing a higher VMC speed. This is normally accomplished during static VMC tests.
(3) Straight Flight. Straight flight is maintaining a constant heading. Section
23.149(a) requires the pilot to maintain straight flight (constant heading). This can be accomplished either with wings level or, at the option of the applicant, with up to five degrees of bank toward the operating engine. Normally, 2-3 de grees of bank allows the airplane to attain zero sideslip so that at five degrees of bank, the beneficial effects of directional stability to counter the yaw produced by asymmetric thrust can be utilized.
(4) Control Forces. The rudder and aileron control force limits may not exceed those specified in § 23.143.
(5) Deicer Boots, Antennas, and other External Equipment. The installation of deicer boots, antennas, and other external gear could change the VMC speed significantly. Re evaluation of the VMC speed should be considered when these installations are made. Refer to AC 23.l4l9-2 if a "flight into icing" approval is being sought.
(6) Variable VMC. For commuter category airplanes, a V MC that varies with altitude and temperature is a permissible condition for use in determining § 23.53 takeoff speeds, provided that the AFM does not show a VR below the red radial line speed required by § 23.1545(b)(6).
(7) Autofeather Annunciations. If autofeather is installed, there should be annunciations to advise of the status. This will include at least green advisory any time the system is armed. For some airplanes, the autofeather system will be identified as a critical system. This could be because VMC has been determined with an operative autofeather system or because commuter category takeoff conditions were predicated on an operative autofeather system. For such installations, additional annunciations may be necessary to ensure that the system is armed and that malfunctions are immediately recognized. This could include caution/warning/advisory annunciations as follows:
(i) Caution or warning, if autofeather switch is not armed.
(ii) Caution or advisory if the autofeather is armed then is subsequently disarmed because of a system malfunction.
(a) All annunciations should be evaluated to verify that they can be easily and quickly recognized. For critical systems, the AFM limitations should require both a satisfactory preflight check and an armed autofeather for takeoff and landing.
c. Procedures.
(1) Configuration. Prior to conducting V MC tests, rudder and aileron control travels should be set to the minimum allowable production travels. Rudder and aileron control cable tensions should be adjusted to the minimum value for use in service. The critical loading for VMC testing is generally minimum weight and maximum aft c.g.; however, each airplane design should be evaluated independently to be assured that tests are conducted under the critical loading conditions. Variable aft c.g. limits as a function of weight, tip tanks, and so forth, can cause the critical loading condition to vary from one airplane to another.
(2) Power. An airplane with a sea-level engine will normally not be able to produce rated takeoff power at the higher test altitudes. Under these circumstances, VMC should be determined at several power settings and a plot of VMC versus power will allow extrapolation to determine VMC at maximum takeoff power. Refer to paragraph c(6) for a further explanation of extrapolation methods. If tests are conducted at less than approximately 3,000 feet density altitude, no corrections to VMC are normally necessary. If tests are conducted above 3,000 feet density altitude, then additional tests should be conducted to allow extrapolation to sea level thrust. Because propeller thrust decr eases with increasing true airspeed, VMC will increase with decreasing altitude and temperature, even at constant power. The results of testing are used to predict the VMC for a maximum takeoff power condition at sea level unless, because of turbocharging or other reasons, some higher altitude prevails as the overall highest VMC value.
(3) Engine Controls. All engine controls have to st ay in the recommended takeoff or approach position, as appropriate, throughout the whole procedure.
(4) Flap Settings. An applicant may want to sp ecify more than one takeoff or landing flap setting, as appropriate, that would require VMC investigation at each flap setting.
(5) Stalls. Extreme caution should be exercised during VMC determination due to the necessity of operating with asymmetric power, full rudder and aileron at speeds near the aerodynamic stall. In the event of inadvertent entry into a stall, the pilot should immediately reduce the pitch attitude, reduce power on the operating engine(s), and return rudder and aileron controls to neutral to preclude possible entry into a spin.
(6) Static Minimum Control Speed. The test pilot should select test altitude based on the capability to develop takeoff power and consistent with safe practices. It will be necessary to determine which engine is critical to the VMC maneuver by conducting static tests with first one then the other engine inoperative to discover which one produces the higher VMC. Power should be set to the maximum available for the ambient condition. If possible, test weights should be light enough to identify the limits of directional control without stalling or being in prestall buffet. For each te st altitude condition, the following should be accomplished:
(i) Flaps and Gear. For the takeoff conditions, the gear should be retracted and the flaps in the takeoff position(s). For the landing conditions, the gear should be extended and the flaps in the landing position(s).
(ii) Trim. The airplane should be trimmed to the settings associated with normal symmetrical power for takeoff or approach (as appropriate) with all engines operating, as indicated.
(iii) Power. Set takeoff power on one engi ne and render the other engine inoperative. The propeller on the inoperative en gine should be windmilling, or in the condition resulting from the availability of automatic feathering or other devices.
(iv) Controls. Gradually reduce airspeed until it is no longer possible to prevent heading changes with maximum use of the directional and near maximum use of the lateral controls, or the limit control forces have been reached. No changes in lateral or directional trim should be accomplished during the speed reduction. Usually the five-degree bank option will be used (refer to paragraph b(3)) to maintain straight flight. A yaw string may be used to assist the test pilot in attaining zero sideslip (or minimum sideslip).
(v) Critical Engine. Repeat steps (a) through (d) to identify which inoperative engine results in the highest minimum control speed. [If an autofeather system is installed and static VMC was determined with the propeller feathered, repeat steps (a) through (d) with the critical engine inoperative and with the propeller windmilling.]
(7) Extrapolation to Sea Level. The only VMC test data that can be extrapolated reliably are static VMC data, where most of the variables can be carefully controlled to a constant value. Because VMC data are typically collected in ambient conditions less critical than sea level standard day, extrapolation is nearly always necessary. Therefore, the usual way to establish an AFM VMC is to extrapolate static VMC data. When V MC is determined for an airplane with an automatically feathered propeller, special techniques are required. Appendix 3 shows one method for extrapolating static VMC from test conditions to sea level standard day.
(8) Dynamic Minimum Control Speed. After determining the critical engine static VMC, and at some speed above static VMC, make a series of engine cuts (using the mixture control or idle cutoff control) dynamically while gradually working speed back toward the static speed. While maintaining this speed after a dynamic engine cut, the pilot should be able to control the airplane and maintain straight flight without reducing power on the operating engine.
During recovery, the airplane should not assume any dangerous attitude nor should the heading change more than 20 degrees when a pilot responds to the critical engine failure with normal 73
skill, strength, and alertness. The climb angle with all engines operating is high, and continued control following an engine failure involves the ability to lower the nose quickly and sufficiently to regain the initial stabilized speed. The dynamic VMC demonstration will normally serve as verification that the numbers obtained statically are valid. If, in fact, the dynamic case is more critical, then the extrapolated static VMC value should be increased by that increment.
Frequently, the dynamic VMC demonstration will indicate a lower VMC than is obtained from static runs. This may be because the inoperative engine, during spooldown, may provide net thrust or control force peaks that exceed limit values for a short period and go undetected, or, due to high yaw and pitch angles and rates, the indicated airspeed values are erroneous. Because of the multi-variable nature of the dynamic VMC demonstration, the AFM VMC value should represent the highest of the static or dynamic VMC test data, corrected to critical conditions.
Specifically in test conditions with a high thrust/weight ratio, a modified procedure may be applied to avoid extreme pitch attitudes. In this case, decelerate to below VMC, all engines, accelerate with two x maximum takeoff power (MTOP) to a representative climb pitch attitude, and cut the critical engine at static VMC (verify in advance that VMC is acceptably above actual stall speed).
(9) Repeatability. Once determined, and if the dynamic VMC seems to be the critical one, the dynamic VMC should be verified by running a series of tests to determine that the speed is repeatable.
(10) AFM Minimum Control Speed Value. V MC is usually observed at either several different power settings or altitudes, or both. Sufficient test data should be obtained such that the VMC for the highest power and sea level density conditions may be determined. The VMC resulting from this extrapolation to sea level is the one entered into the AFM and marked on the airspeed indicator. If this V MC is determined with an autofeather system, the AFM required equipment list, as well as the Kind of Equipment List (KOEL), should list autofeather as a required item and the AFM would normally state the VMC with the autofeather system inoperative (propeller windmilling) in the procedures section. The procedures section should also require the autofeather to be armed (if applicable) during takeoff and landing.
(12) Safe, Intentional, One-Engine Inoperative Speed, VSSE. Reserved.
5. § 23.151 Acrobatic Maneuvers.
a. Explanation. This regulation requires each maneuver to be evaluated and safe entry speeds established. Section 23.1567( c), which is associated with this requirement, imposes a requirement for a placard that gives entry airspeeds and approved maneuvers. If inverted flight is prohibited, the placard should so state.
b. Procedures. The applicant should fly each maneuver for which approval is sought.
The FAA test pilot should then evaluate those maneuvers considered most critical.
c. Data Acquisition. A recently calibrated airspeed system, airspeed indicator, accelerometer, and tachometer should be provided by the applicant for the test airplane. The following should be recorded:
(1) Load factor;
(2) Entry airspeeds;
(3) Maximum airspeeds; and 74
(4) Maximum r.p.m.
6. § 23.153 Control During Landings.
a. Explanation.
(1) Purpose. The purpose of this requirement is to ensure that airplanes do not encounter excessive control forces when approaching at a speed of five knots lower than normal landing approach speed. Also, a safe landing is re quired. Safe is considered to include having sufficient flare capability to overcome any excessive sink rate that may develop.
(2) Landing Requirements. Section 23.75 is a companion requirement and normally tests to determine compliance would be accomplished at the same time.
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.
7. § 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 in flight 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 qualitativ e 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 shown in the table below should be investigated in the cruise configuration, that is, flaps up and gear up (if retractable):
Table 7 - Procedures Condition Power Wings Level TrimSpeed Altitude 1 Refer to Note Trimmed (but not to exceed VNE or VMO/MMO) Altitude for highest dynamic pressure (q) 3 Refer to Note VO Low 4 Refer to Note VO Highest attainable approved altitude Note: 75 percent maximum continuous power or maximum power (reciprocating engine) or maximum cruise power (turbine).
(1) Compliance may be demonstrated by measuring the normal acceleration achieved with the limiting stick force (50 pounds for wheel controls, 35 pounds 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) Weight/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.
Note: The test data should be presented in stick force versus g plots. Figure 17 below shows a sample plot. Test results should be compared to the requirements of § 23.155(a).
~ ,~ -.... .... .... .... ....
I,' ,,.. ....
I,' ....
- ,..
/ / , ;, / >- / / ,I ~ / ,,, J , 11/16/2011 AC 23-8C Figure 17 - Stick Force Per G 50 40 30 LBS PULL 20 10 0 CAS= ________
c.g.= ________
0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 NORMAL ACCELERATION - G's
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 elevator cont rol force should increase progressively with increasing load factor.
(1) 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:
(i) The intensity of buffet provides a strong and effective deterrent to further increase of load factor;
(ii) Further increase of load factor requires an elevator control force in excess of 150 pounds for a wheel control or 125 pounds for a stick control or is impossible because of the limitations of the control system; or
(iii) The positive limit maneuvering load factor is achieved.
8. § 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.
Section 5. Trim
1. § 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. If appropriate, lateral baggage loading and fuel asymmetry should be considered in this evaluation. While ground adjustable tabs are acceptable, they are typically installed on simple single engine airplanes and aren’t effective for addressing control forces with significant asymmetry loadings. Applicants should either provide a way to set the ground adjustable tab for specific loadings or install an air adjustable tab. For multiengine airplanes, directional trim is required for a range of speeds.
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.
Section 6. Stability
1. § 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 78
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 fo rces should increase with departure from the trim speed up to the speed limits specified in § 23.175 or up to the 40 pound force limit specified in § 23.173. The stick forc e 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 18 below shows an exam ple of cruise configuration:
PULL (+) STICK FORCE PUSH (-) FRSR +> 40 KNOTS OR 15%VTRIM FRSR +> 40 KNOTS OR 15%VTRIM VTRIM j 11/16/2011 AC 23-8C Figure 18 - Static Longitudinal Stability Data
(i) Speed Range = Greater of + 40 knots (kts) or 15 percent VTRIM + free return speed range (FRSR)
(ii) 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, refer to figures 19 and 20. The pull forces can decrease in magn itude 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.
(iii) Instrumented force measurements are required if there is any uncertainty in the qualitative assessment of the force gradients.
Figure 19 - Low Speed Instabilities; Normal, Utility, and Acrobatic Airplanes FRSR +> 40 KNOTS OR 15% V TRIM FRSR +> 40 KNOTS OR 15% V TRIM PULL (+)
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