(vi) Cowl flap position may be obtained from a calibrated indicator or a measured position.
(vii) Wing flap position may be obtained from a calibrated indicator or a measured position.
(viii) Slope of the runway can be obtained from the official runway survey or other suitable data obtained using accepted survey practices.
(ix) Direction of the landing run will be the direction of the runway used or an accurate compass indication.
(x) The wind direction and velocity should be obtained with an accurate compass and a calibrated anemometer. Wind data obtained from airport control towers should not be used.
14. § 23.77 Balked Landing Climb.
a. Explanation (Normal, Utility, and Acrobatic Category). Reciprocating engine airplanes with a MTOW of 6,000 pounds (2730 kg) or less.
(1) Purpose. The configuration that is specified for this climb requirement ordinarily is used in the final stages of an approach for landing. The objective of requiring the prescribed climb capability is to ensure that the descent may readily be arrested, and that the airplane will be able to "go around" for another attempt at landing, in the event conditions beyond control of the pilot make such action advisable or necessary.
(2) Flap Retraction. As an alternative to having the flaps in the landing position, compliance with the balked landing climb requirement may be demonstrated with flaps in the retracted position, provided the flaps are capable of being retracted in two seconds or less. Also, provided the airplane's flight characteristics during flap retraction satisfy the constraints imposed by the regulation; that is, flaps must be retracted with safety, without loss of altitude, without sudden change in angle-of-attack, and without need for exceptional piloting skill. Evaluation should include satisfactory demonstration of ability to promptly arrest the descent by application of takeoff power in conjunction with rapid retraction of the flaps during final approach to landing.
(3) Flaps That Will Not Fully Retract in Two Seconds. If the flaps will not fully retract in two seconds, the climb available with the flap position at the end of two seconds may be used and the position noted in the AFM. Other considerations should include flight characteristics, ease of operation and reliability. If the flap is not mechanical, the flap mechanism should be reliable in order to receive credit for a partially retracted flap.
b. Procedures. Climb performance tests are conducted to establish compliance with the prescribed climb requirement and for inclusion in the AFM. The procedures outlined under § 23.65 are equally applicable to the balked landing climb, except that the cooling and other 63
considerations that recommend exploration of a speed range by conducting sawtooth climbs do not apply to the balked landing climb. In lieu of sawtooth climbs, the balked landing climb performance may be established as the average of not less than three continuous run pairs at the climb speed selected by the applicant.
c. Explanation (Normal, Utility and Acrobatic Aircraft with MTOW greater than 6,000 pounds (2730 kg) and Turbine Engine Aircraft and Commuter Category Aircraft).
Sections 23.77(b)(1) and 23.77(c)(1) state that the engines are to be set at the power or thrust that is available eight seconds after initiation of movement of the power controls from minimum flight idle to the takeoff position. The procedures given are for the determination of this maximum power for showing compliance with the climb requirements of § 23.77.
d. Procedures. (Aircraft with a MTOW greater than 6,000 pounds and turbine-engine aircraft.)
(1) Engine Trim. Trim engines to the minimum idle speed/power to be defined in the airplane maintenance manual.
(2) Engine Power Tests. Engine power tests should be conducted at the most adverse landing elevation and temperature condition, or the range of landing altitude and temperature conditions if the most adverse cannot be readily determined.
(i) In the critical air bleed configuration, if applicable, stabilize the airplane in level flight with symmetrical power on all engines, landing gear down, flaps in the landing position, at a speed of VREF at an altitude sufficiently above the selected test altitude so that time for descent to the test altitude with all throttles closed will result in minimum flight-idle power at test altitude.
(ii) Retard throttles to flight idle and descend at VREF, as defined in § 23.73, to approximately the test altitude. When power has stabilized, advance throttle(s) in less than one second to obtain takeoff power.
(iii) The power that is available eight seconds after the initiation of movement of the power controls from the minimum flight idle position is the maximum permitted for showing compliance with the landing climb of § 23.77 for each of the bleed combinations tested.
(iv) If AFM performance is presented so there is no accountability for various bleed conditions, the power obtained with the most critical air bleed should be used for landing climb performance for all operations, including the effects of anti-ice bleed.
e. Data Acquisition and Reduction. The information presented under § 23.65 of the regulation applies to the balked landing climb.
Section 3. Flight Characteristics of 14 CFR § 23.141
1. § 23.141 General
a. Explanation.
(1) Minimum Flight Characteristics. The purpose of these requirements is to specify minimum flight characteristics that are considered essential to safety for any airplane. This section deals primarily with controllability and maneuverability. A flight characteristic is an attribute, a quality, or a feature of the fundamental nature of the airplane that is assumed to exist because the airplane behaves in flight in a certain consistent manner when the controls are placed in certain positions or are manipulated in a certain manner. In some cases, measurements of forces, control surface positions, or acceleration in pitch, roll, and yaw may be made to support a decision but, normally, it will be a pass/fail judgement by the Civil Aviation Authority’s/FAA test pilot.
(2) Exceptional Skills. The phrase "exceptional piloting skill, alertness, or strength," is used repeatedly throughout the regulations and requires highly qualitative judgments on the part of the test pilot. The judgments should be based on the pilot's estimate of the skill and experience of the pilots who normally fly the type of airplane under consideration (that is, private pilot, commercial pilot, or airline transport pilot skill levels). Exceptional alertness or strength requires additional judgment factors when the control forces are deemed marginal or when a condition exists that requires rapid recognition and reaction to be coped with successfully.
b. Procedures. None.
Section 4. Controllability and Maneuverability
1. § 23.143 General.
a. Explanation.
(1) Temporary Control Forces. Temporary application, as specified in the table in § 23.143(c)(1), may be defined as the period of time necessary to perform the necessary pilot motions to relieve the forces, such as trimming or reducing power. The values in the table under § 23.143 of part 23 are maximums. There may be circumstances where a lower force is required for safety. If it is found that a lower force is nece ssary for safety, then that lower force should be established under § 21.21(b)(2).
(2) Prolonged Control Forces. Prolonged application would be for some condition that could not be trimmed out, such as a forward c.g. landing. The time of application would be for the final approach only if the airplane could be flown in trim to that point. Section 23.689 includes discussion concerning a failed cable. .
(3) Controllability. Controllability is the ability of the pilot, through a proper manipulation of the controls, to establish and maintain or alter the attitude of the airplane with respect to its flight path. It is intended in the design of the airplane that it be possible to "control" the attitude about each of the three axes, the longitudinal, the lateral, and the vertical axes. Angular displacements about the longitudi nal axis are called "roll." Those about the lateral axis are called "pitch" and those about the directional axis are called "yaw."
Controllability should be defined as "satisfactory" or "unsatisfactory." Unsatisfactory controllability would exist if the test pilot finds the controllability to be so inadequate that a 65
dangerous condition might easily occur and is unacceptable as a showing of compliance with the regulations.
(4) Maneuverability. Maneuverability is the ability of the pilot, through a proper manipulation of the controls, to alter the direction of the flight path of the airplane. In order to accomplish this, the airplane should be controllable, since a change about one of the axes is necessary to change a direction of flight. Any change in the direction of flight involves an acceleration normal to the flight path. Maneuverability is so closely related to controllability as to be inseparable in any real motion of the airplane. It is also similarly largely qualitative in its nature and should be treated in the same manner as has been suggested for controllability above.
(5) Spring Devices. If a spring device is insta lled in the control system, § 23.687 requires that the airplane not have any unsafe flight characteristics without the use of the spring device, unless the reliability of the device can be established by tests simulating service conditions. The airplane does not have to meet the handling, stability, and control requirements with a failed spring device; however, the airplane has to demonstrate that it is easily controllable by the pilot to continue flight to a safe landing with the failed spring(s).
(6) Powered Trim Without Manual Backup. Most of the part 23 airplanes built since the 1990s use some type of powered trim and have eliminated the manual trim wheel that was traditionally available to the pilot. This configuration makes the failure of the trim system more critical than in the past because pilots cannot change the trim position after the failure. If the failure of the trim system occurs during a slower flight regime such as climb, the control forces to return and land should be reasonable. If, on the other hand, the failure occurs during highspeed cruise in a fast airplane, the control forces when slowing down to land may exceed the prolonged (and maybe even the temporary) force limits in § 23.143. Applicants should account for this possibility and consider mitigation strategies to include in the AFM on how to handle this failure condition.
b. Procedures.
(1) Landing. Using the AFM recommended approach/landing speeds and power settings, determine that airplane controllability is satisfactory with the wing flaps extended and retracted. These tests should be accomplished at the critical weight/c.g. combination within the allowable landing range. For turboprop airplanes, the engine fuel control should be adjusted to the minimum flight-idle fuel flow permitted on airplanes in service unless it is shown that the range of adjustment permitted on airplanes in service has no measurable effect on flight-idle sink rate.
(2) Other Flight Conditions. Controllability and mane uverability procedures for other flight conditions, such as takeoff and VMC, are covered in their respective sections.
(3) Lateral imbalance. Lateral imbalance flight evaluations should be conducted on all airplanes configured such that lateral trim and controllability may be affected. The following configurations should be considered and evaluated as appropriate:
(i) Takeoff - All engine, one-engine-inoperative (multiengine airplanes), VMC, and crosswind operations.
(ii) Enroute - All engine, one-engine-inoperative (multiengine airplanes), and autopilot coupled operations.
(iii) Approach and Landing - All engine, one-engine-inoperative (multiengine airplanes), VMC (where applicable), crosswind, and autopilot coupled operations.
(a) As a result of flight tests, appropriate lateral imbalance limitations and procedures should be developed. Different valu es of fuel imbalance for the various flight configurations may be required. Imbalance limits, if any, should be included in the AFM.
c. Data Acquisition and Reduction. A qualitative determination by the test pilot will usually suffice unless the control force limits are considered marginal. In this case, force gauges are used to measure the forces on each affected control while flying through the required maneuvers.
2. § 23.145 Longitudinal Control.
a. Explanation.
(1) Elevator Power. This regulation requires a seri es of maneuvers to demonstrate the longitudinal controllability during pushovers from low speed, flap extension and retraction, and during speed and power variations. The prime determinations to be made by the test pilot are whether or not there is sufficient elevator power to allow pitching the nose downward from a minimum speed condition and to assure that the required maneuvers can be performed without the resulting temporary forces becoming excessive.
(2) Speeds Below Trim Speeds. The phrase, "speeds below the trim speed," as used in § 23.145(a), means speeds down to VS1.
(3) Wing Flaps. If gated or detented flap pos itions are provided, refer to the requirements in § 23.697(c). (An acceptable gate or detent is one that allows the pilot to readily determine the flap position without seriously detracting from other piloting duties under any flight condition, day or night.)
(4) Loss of Primary Control Systems. Section 23.145(e) is intended to cover a condition where a pilot has sustained some failure in the primary longitudinal control system of the airplane (for some multiengine airplanes, also loss of the directional control system) and is required to land using the power and trim system without the primary control. It is not intended that this test be demonstrated to an actual landing; however, a demonstration may be performed using manipulation of trim and power to a landing, if desired. Section 23.145(e) is the flight test to demonstrate compliance with the requirement that specifies a failure of the primary control system.
(5) Analysis of System. An analysis of the control system should be completed before conducting the loss of primary control system test. On some airplanes, the required single longitudinal control system failure could result in loss of both the downspring and the primary longitudinal control system. If this failure occurred on an airplane utilizing an extremely large 67
downspring, the loss of the downspring may result in a nose-up pitching moment at aft c.g. that could not be adequately countered by the basic pitch trim system.
b. Procedures. The wording of the regulation sufficiently describes the maneuvers required to show compliance. The selection of altitudes, weights, and c.g. positions to be flight tested by the FAA will depend on a study of the applicant's flight test report. Normally, the following combinations are checked during the certification tests:
(1) Altitude. Select a nominal altitude considering power available. Longitudinal control would not be expected to change much with typical part 23 altitude capabilities, so repeating tests near maximum altitude should only be needed when controllability is marginal and/or the airplane has high altitude capability like some of the part 23 jets.
(2) Weight. Maximum gross weight for all tests, except where otherwise described in paragraph (3) below.
(3) C.G. For conventional configurations, § 23.1 45(a), most aft c.g. and most aft c.g.
approved for any weight; § 23.145 paragraphs (b) 1 through (b) 6, most forward and most aft
c.g.; § 23.145(c), most forward c.g.; § 23.145(d), most forward c.g. and most forward c.g.
approved for any weight; and § 23.145(e), both the forward and aft c.g. locations. Section 23.145(e) is sometimes more difficult to achieve at the aft c.g. than the forward limit, particularly if the airplane exhibits neutral to divergent phugoid tendencies.
(4) Power or Configuration. Pitching moments resulting from power or configuration changes should be evaluated under all conditions necessary to determine the most critical demonstration configuration.
c. Data Acquisition. No special instrumentation is required. The exception to this would be the l0-pound force in § 23.145(d), which should be measured with a force gauge. All longitudinal forces should be measured if the forces are considered marginal or excessive.
3. § 23.147 Directional and Lateral Control.
a. Explanation.
(1) Yawed Flight. Section 23.147(a) is intended as an investigation for dangerous characteristics during sideslip, which may result from blocked airflow over the vertical stabilizer and rudder. Rudder lock and possible loss of directional control are exam ples of the kinds of characteristics the test is aimed at uncovering. Section 23.177 also addresses rudder lock.
Compliance may be demonstrated if the rudder stop is reached prior to achieving either l5 degrees of heading change or the l50-pound force limit, providing there are no dangerous characteristics. The control stop serves more effectively than the l50-pound force to limit the pilot's ability to induce a yaw beyond that demonstrated as acceptable.
(2) Controllability following sudden engine failure. Section 23.147(b) requires a demonstration of controllability following sudden engine failure during en route climb.
b. Procedures.
(1) Yawed Flight. The airplane configurations to be tested according to § 23.147(a) are:
(i) One engine inoperative and its propeller in the minimum drag position.
(ii) The remaining engines at not more than maximum continuous power.
(iii) The rearmost allowable c.g.
(iv) The landing gear: both retracted and extended
(v) The flaps retracted.
(vi) Most critical weight.
(vii) Airplane trimmed in the test condition, if possible.
(2) Controllability Following Sudden Engine Failure. In complying with the testing required by § 23.147(b), from an initial condition of straight flight with wings level, zero sideslip and in trim, simulate a sudden and complete failure of the critical engine. To allow for an appropriate delay, no action should be taken to recover the airplane for two seconds following the first indication of engine failure. The recovery action should not involve movement of the engine, propeller or trimming controls. At no time should the bank angle exceed 45 degrees or excessive yaw be developed. The evaluation of da ngerous attitudes and characteristics should be based on each particular airplane characteristic and the flight test pilot's evaluation.
(i) The method used to simulate engine failure should be:
(a) For a reciprocating engine, closure of the mixture control; or
(b) For a turbine engine, termination of the fuel supply by the means which results in the fastest loss of engine power or thrust. Engine shut-off procedures would normally be sufficient.
c. Loss of Primary Control Systems. (Also, refer to AC 23-17.)
(1) Explanation. Section 23.147(c) is intended to cover a condition where a pilot has sustained some failure in the primary lateral control system of the airplane. If a single failure in the primary lateral control system could also cause the loss of additional control, then the loss of the additional controls must be considered. It must be shown that, with the loss of the primary lateral control, the airplane is safely controllable in all configurations and could be landed without exceeding the operational and structural limitations of the airplane. The rule calls out any engine configuration, speed, or altitude. This rule is interpreted to mean safely controllable such that the pilot can transition into a smaller, limited flight and configuration envelope, if necessary, for continued safe flight and landing. It is not intended that this test be demonstrated to an actual landing; however, a demonstration may be performed using manipulations of either lateral trim or sideslip, or both, generated by the rudder and differential power if available to a landing. Section 23.147(c) is the flight test to de monstrate compliance with the requirement that specifies a failure in the primary lateral control system. This failure implies a disconnection on the primary control system such that the ailerons are free to float and the lateral trim (if installed) is operational.
(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