necessary to assure that the airplane is firmly on the surface. Associated procedures should be included in the AFM.
(v) Reverse Thrust for Multiengine Airplanes. In the approval of reverse thrust for turboprop airplanes, due consideration should be given for thrust settings allowed, the number of operating engines, and control of the aircraft with one engine inoperative. If landing distance depends on the operation of any engine and if the landing distance would be noticeably increased (2 percent has been found acceptable) when a landing is made with that engine inoperative, the landing distance should be determined with that engine inoperative unless the use of compensating means (such as reverse thrust on the operating engine) will result in a landing distance not more than that with each engine operating (this assumes that there are no other changes in configuration, for example, flap setting associated with one engine inoperative, that will cause an increase in landing distance). In effecting thrust changes following touchdown, allowance should be made for any time delays that reasonably may be expected in service, or which may be necessary to assure that the airplane is firmly on the surface. See subparagraph b(2) for commuter category time delays. Associated procedures should be included in the AFM.
(vi) Reverse Thrust for Single-Engine Airplanes. In effecting thrust changes following touchdown, allowance should be made for any time delays that reasonably may be expected in service, or which may be necessary to assure that the airplane is firmly on the surface. Associated procedures should be included in the AFM.
(6) Balked Landing Transition. For the power conditions selected for the landing demonstration (except one engine inoperative), and other steady state conditions of speed and rate of sink that are established during the landing approach, it should be possible, at the 50-foot point, to make a satisfactory transition to the balked landing climb requirement of § 23.77 using average piloting skill without encountering any unsafe conditions.
(7) Expansion of Landing Data for a Range of Airport Elevations. When the basic landing tests are accomplished between sea level and approximately 3,000 feet, the maximum allowable extrapolation limits are 6,000 feet above and 3,000 feet below the test field elevation. If you desire to extrapolate beyond these limits, one of two procedures may be employed. These procedures are given in paragraph 19c(3) (c).
b. Procedures.
(1) Technique. The landing approach should be stabilized on target speed, power, and the airplane in the landing configuration prior to reaching the 50-foot height to assure stabilized conditions when the airplane passes through the reference height. The engine fuel control should be adjusted to the maximum flight-idle fuel flow permitted on airplanes in service unless it is shown that the range of adjustment has no effect on landing distance. A smooth flare should be made to the touchdown point. The landing roll should be as straight as possible and the airplane brought to a complete stop (or 3 knots for seaplanes) for each landing test. Normal pilot reaction times should be used for power reduction, brake application, and use of other drag/deceleration devices. See subparagraph b(2) for commuter category time delays. These reaction times should be established by a deliberate application of appropriate controls as would be used by a normal pilot in service. They should not represent the minimum times associated with the reactions of a highly trained test pilot.
(2) Commuter Category Time Delays.
(i) The time delays shown in figure 27-1 should be used.
(ii) For approved automatic deceleration devices (for example, autospoilers, etc.) for which performance credit is sought for AFM data expansion, established times determined during certification testing may be used without the application of the one-second minimum time delay required in the appropriate segment above.
(3) Applicant's Procedures. The procedures to be followed should be those recommended by the applicant.
TOUCH DOWN PILOT
ACTUATION
OF FIRST
DECELERATION DEVICE PILOT
ACTUATION
OF SECOND
DECELERATION
DEVICE STOP 12 Transition from touchdown to full braking configuartion Full braking configuration to stop
c - This segment represents the flight test measured average time from touchdown to pilot actuation of the first deceleration device. For AFM data expansion, use one second or the test time, whichever is longer.
d - This segment represents the flight test measured average test time from pilot actuation of the first deceleration device to pilot actuation of the second deceleration device. For AFM data expansion, see item c above.
Segment d is repeated until pilot actuation of all deceleration devices has been completed and the airplane is in the full braking configuration.
FIGURE 27-1 - LANDING TIME DELAYS
(4) Number of Landings. At least six landings should be conducted on the same wheels, tires, and brakes to establish the proper functioning required by § 21.35(b).
(5) Winds. Wind velocity and direction should be measured adjacent to the runway during the time interval of each test run. See paragraph 6a(5) of this AC for wind velocity and direction tolerances.
(6) Weight. Landing tests should be conducted at maximum landing weight.
(7) Approach Angles Greater than Three Degrees. If the applicant chooses an approach angle greater than three degrees, landing distances that result from utilizing a three-degree approach angle should be determined and published in the AFM to enable operators to comply with related operational rules.
c. Data Acquisition.
(1) The data to be recorded for landing distance tests are:
(i) Vertical and horizontal path of the airplane relative to the runway. Two methods that have been used are runway observers and time histories. Sink rate at touchdown and descent gradients may be computed from time histories.
(ii) Pressure altitude.
(iii) Ambient air temperature.
(iv) Airplane weight (fuel used or time since engine start).
(v) Engine power or thrust data.
(vi) Cowl flap position.
(vii) Wing flap position.
(vii) Runway slope.
(ix) Direction of landing run.
(x) Wind direction and velocity at a height of 6 feet adjacent to the runway near the touchdown point.
(xi) Landing procedures noted for inclusion in the AFM.
(2) Means of acquiring the required data are listed below:
(i) Time history data is obtained by use of a takeoff and landing camera, electronic equipment, or a phototheodolite having a known surveyed location. If landing gear loads are a concern, sink rate at touchdown may be computed, or alternately, vertical load factor may be measured by an accelerometer at the c.g.
(ii) Pressure altitude may be obtained with a calibrated sensitive altimeter.
(iii) Ambient air temperature should be obtained with a calibrated temperature sensor.
(iv) The airplane weight may be computed from a known weight at start of test minus the fuel used to the time of test.
(v) Engine power or thrust data may be determined using calibrated airplane powerplant instruments to provide the basic parameters required.
(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.
30. SECTION 23.77 BALKED LANDING CLIMB.
a. Explanation (Normal, Utility, and Acrobatic Category). Reciprocating engine airplanes with a MTOW of 6,000 lbs. (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 2 seconds or less and 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 as a consideration in an equivalent level of safety finding. 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 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 56
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 lbs.
(2730 kg) and Turbine Engine Aircraft and Commuter Category Aircraft). Section
23.77(b)(1) states that the engines are to be set at the power or thrust that is available 8 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 applies to the balked landing climb.
31.-38. RESERVED.
Section 3. FLIGHT CHARACTERISTICS
39. SECTION 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 Authority 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.
(3) Stall Speed Multipliers. For conventional configurations, flying qualities and trim speeds may only be based on the forward c.g. stall speeds.
b. Procedures. None.
40.-44. RESERVED.
Section 4. CONTROLLABILITY AND MANEUVERABILITY
45. SECTION 23.143 GENERAL.
a. Explanation.
(1) Temporary Control Forces. Temporary application, as specified in the table, 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 necessary 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.
(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 directional axes.
Angular displacements about the longitudinal 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 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 installed 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.
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 maneuverability 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.
As a result of flight tests, appropriate lateral imbalance limitations and procedures should be developed. Different values 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.
46. SECTION 23.145 LONGITUDINAL CONTROL.
a. Explanation.
(1) Elevator Power. This regulation requires a series 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 flap positions are provided, see § 23.697. (An acceptable gate 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 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 Authority will depend on a study of the applicant's flight test report. Normally, the following combinations are checked during the certification tests:
(1) Altitude. A low altitude and an altitude near the maximum altitude capability of the airplane. A high altitude may not be needed for normally aspirated engine airplanes.
(2) Weight. Maximum gross weight for all tests, except where otherwise described in subparagraph (3) below.
(3) C.G. For conventional configurations, § 23.145(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.
47. SECTION 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 examples 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:
• 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 first indication of engine failure. The recovery action should not involve movement of the engine, propeller or trimming controls. At no time until the completion of the maneuver should the bank angle exceed 45 degrees or excessive yaw be developed. The evaluation of dangerous attitudes and characteristics should be based on each particular airplane characteristic and the flight test pilot's evaluation.
The method used to simulate engine failure should be:
(i) For a reciprocating engine, closure of the mixture control; or
(ii) 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, see 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. 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 demonstrate 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 for the most critical case. If compliance with § 23.147(c) can 62
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. A low altitude and an altitude near the maximum capability of the airplane.
The high altitude test is to determine controllability with decreased Dutch roll damping.
(ii) Weight. Maximum gross weight for all tests, except where otherwise described in subparagraph (3) below.
(iii) C.G. For conventional configuration, § 23.147(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.
48. SECTION 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. VMC 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. VMC 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 VMC is directly related to the control surface travel available. Normally, VMC is based on available rudder travel but may, for some airplanes, be based on lateral control travel. 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 with 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 63
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 VMC 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 VMC 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 5 degrees of bank toward the operating engine. Normally, 2-3 degrees of bank allows the airplane to attain zero sideslip so that at 5 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. See AC 23.l4l9-2 if a "flight into icing" approval is being sought.
(6) Variable VMC. For commuter category airplanes, a VMC 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.