AC 27.1B, Chg4 9/17/2009 AC 27.175A. § 27.175 (Amendment 27-44) DEMONSTRATION OF STATIC LONGITUDINAL STABILITY.
a. Explanation. Amendment 27-44 reduces the speed range for the climb and cruise demonstration points of §§ 27.175(a) and 27.175(b), respectively. A new paragraph (c) was added to require an additional cruise demonstration point in order to compensate for the change in reduced speed range in paragraph (b). Additionally, for autorotation, two typically used trim points are required in place of the current requirement. The requirement for the hover demonstration was eliminated for the reasons given in AC 27.173 (Amendment 27-44).
All the policy material pertaining to this section remains in effect with the following changes:
(1) This rule incorporates the specific flight requirements for demonstration of static longitudinal stability. Specific loadings, configurations, power levels, and speed ranges are stated for conditions of climb, cruise, VNE, and autorotation.
(2) Some rotorcraft in forward flight experience significant changes in engine power with changes in airspeed even though collective and throttle controls are held fixed and altitude remains relatively constant. For these cases, the guidance in § 27.173, which states that throttle and collective pitch must be held constant, is appropriate for administration of this rule, and the specified powers in § 27.175 should be considered as power established at initial trim conditions. This will result in slightly higher or lower power readings at “off trim” conditions. Collective and throttle controls are held constant when obtaining test data.
(3) The effects of rotor RPM on autorotative static stability should be determined and positive stability demonstrated for the most critical RPM. For Category A rotorcraft this requirement may be satisfied at a nominal RPM value. RPM values can be expected to change as airspeed is varied from the “trimmed” condition. The manufacturer’s recommended autorotation airspeed is ordinarily used for trim.
b. Procedures.
All the policy material pertaining to this section remains in effect with the following changes:
(1) Instrumentation.
(i) Sensitive control position instrumentation is mandatory. Engine power parameters should be recorded at trim. For testing of minor modifications or when using a “before and after” method, a tape measure or a stick plotting board may be utilized. A stick plotting board consists of a level surface with a clean sheet of paper on it and attached to the cockpit or seat structure. The installation must not interfere when the flight controls are fully displaced. A recording pencil is attached to the cyclic control Page B – 60.2
by an offsetting arm in such a manner that it can be pushed down on the board to record relative cyclic position at key times during test maneuvers. The Figure AC
27.175A-1 plot is a typical presentation of longitudinal static stability.
(ii) Other necessary parameters include pitch attitude, pressure altitude, ambient temperature, and indicated airspeed.
(2) Ambient Conditions. Smooth air is necessary for stability testing.
(3) Loading. Aft center of gravity (CG) is ordinarily critical for longitudinal stability testing, although high-speed flight should be checked at full forward CG and maximum weight. At aft CG, light or heavy weight conditions can be critical. The manufacturer’s flight data should be reviewed to determine critical loading conditions.
(4) Conducting The Test.
(i) The rotorcraft should be established in the desired configuration and flight condition (climb, cruise, VNE, autorotation) with the required power and rotor speed at the trim airspeed. The collective stick should be fixed in that position: usually by applying sufficient friction to ensure that it is not inadvertently moved. For autorotative tests, a rotor speed should be selected so that the variations in rotor speed as airspeed and altitude change do not exceed the allowable limits. This point is recorded as the trim point. Airspeed is then increased or decreased in about 5- knot increments, stabilizing on each speed and recording the data. At least two points on each side of the trim speed should be taken.
(ii) The cruise test should be conducted by varying airspeed around the desired altitude with throttle and collective fixed. This should be accomplished by first determining VH (level flight speed at maximum continuous power) at the test altitude.
Then adjust power to establish a level trimmed condition at VH (or 0.8 VNE if lower). This point is then recorded as the trim point.
(iii) For climb and autorotation tests, conduct fixed collective tests through an altitude band (usually ±2,000 feet). It will probably not be possible to obtain the required data on one pass through the altitude band. If repeated passes are required, a trim point should be taken at the beginning of each pass unless very sensitive collective pitch position information is available in the cockpit.
(iv) If extremely precise results are required, an alternate method of testing can be used to acquire the data at a constant altitude. For cruise and VNE, data can be obtained by alternating airspeeds above and below the trim speed to arrive in the vicinity of the test altitude as the point is recorded. This method results in very precise data because collective and throttle are not moved as airspeed is changed at a constant altitude. A typical sequence of speeds that could produce these results would be: 140
(0.8 VNE) trim speed, 135, 145, 130, and 150.
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AC 27.1B, Chg4 9/17/2009
(v) For rotorcraft with high rates of climb, a series of climbs, each at a different speed, may be required through a given altitude, utilizing sensitive instrumentation to assure collective position is the same for each data point. In autorotation, a similar case arises and a series of descents, each at a different speed, may be required through a given altitude band, using sensitive instrumentation to assure a repeatable collective position.
(vi) Normally tests should be conducted at low, medium, and high altitudes.
See AC 27.45 for guidance on interpolation and extrapolation. High-speed stability has been critical during cold weather testing. Cold weather testing should be accomplished or a conservative approach for advancing blade tip Mach number should be used to limit cold weather VNE to tip Mach number values demonstrated.
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Climb (Vy) Cruise (Vne) Criuse (Vh) Autorotation (Vglide) Autorotation (Vmin RoD) 0 10 20 30 40 50 60 70 80 90 100 Longitudinal Cyclic Position (% from Full Aft) 0 20 40 60 80 100 120 140 160 Calibrated Airspeed (Knots) FIGURE AC 27.175A-1
STATIC LONGITUDINAL STABILITY Page B – 60.5
AC 27.1B, Chg4 3/31/2014 AC 27.177. § 27.177 (Amendment 27-21) STATIC DIRECTIONAL STABILITY.
a. Explanation. This rule requires that positive static directional stability be demonstrated at the trim airspeeds defined in § 27.175. The trim speed for climb is VY and for cruise 0.9VH or 0.9VNE (whichever is less).
b. Procedures.
(1) Tests for static directional stability require instrumentation for pedal position and sideslip angle. To obtain accurate sideslip angle and airspeed information, a “yaw boom” is usually installed for the purpose of mounting a sideslip vane and swiveling airspeed pilot head outside the main rotor downwash region of influence. Special care should be taken to ensure that the yaw boom installation has been verified to be structurally adequate and free of dynamic instabilities for all combinations of airspeed and rotor speed likely to be experienced during the static directional evaluation. For some installations, the instrumentation yaw boom may influence the flying qualities of the rotorcraft itself. Thus, it is advisable to correlate yaw string displacement or slip indicator ball widths of skid with yaw boom sideslip angle, and then repeat a few critical points with the yaw boom removed.
(2) For some rotor system designs, the main and tail rotor flapping angle may be a critical instrumentation requirement for static directional testing. Both main and tail rotor flapping may increase dramatically at high airspeeds with increasing sideslip angle. Therefore, for rotor systems exhibiting this characteristic, flapping should be monitored carefully during the sideslip maneuver to avoid exceeding limitations. Static directional stability is normally defined in terms of pedal displacement required to maintain a steady heading sideslip. A single- rotor rotorcraft flying in coordinated flight will exhibit a small inherent sideslip due to tail rotor thrust and fuselage/main rotor sideforces. This condition is normally taken as trim with the inherent sideslip angle noted. Airspeeds should be the trim values described above. The procedures used to establish and maintain the steady heading sideslip can significantly influence test results. A generally accepted technique follows:
(i) Stabilize at the trim point, and note indicated airspeed.
(ii) Record trim conditions including inherent sideslip. Maintain fixed collective and throttle for the remainder of the maneuver.
(iii) Smoothly apply directional control and coordinate with lateral control to establish the desired sideslip angle. A steady track can best be ensured by maintaining a track over a straight landmark on the ground such as a section line or straight segment of powerline or highway.
(iv) Because drag increases with sideslip, the aircraft will decelerate. The trim airspeed should be maintained by entering a slight dive or decreased rate of climb.
If a boom airspeed system is not used, a “standard” airspeed system will become Page B – 60.6
excessively inaccurate at about 10° of sideslip. Rapidly yawing to the desired sideslip angle and noting and maintaining the new “standard” indicated airspeed may give adequate data. Control positions (directional as a minimum), sideslip angle, rotorspeed, airspeed, rate of descent, amount of ball deflection, and bank angle should be recorded.
The pilot should note the physical sideforce feel experienced. The rule requires that sufficient cues accompany sideslip to alert the pilot when approaching sideslip limits. A minimum of two sideslip data points on each side of the trim point should be obtained to adequately define the slope of the pedal displacement versus sideslip angle relationship.
(v) Static directional stability plots can be expected to differ slightly on either side of the inherent sideslip angle. Positive static directional stability is indicated by increased left pedal displacement for a larger right sideslip and, conversely, increased right pedal for a larger left sideslip angle.
AC 27.177A. § 27.177 (Amendment 27- 44) Static Directional Stability.
a. Explanation. Amendment 27-44 makes an extensive change to the current requirement and provides for a clear definition of the sideslip envelope to be evaluated.
Most rotorcraft exhibit satisfactory quantitative and qualitative directional characteristics except for the first 2-3 degrees either side of trim due to inherent airflow blockage of the vertical fin or tail rotor. This amendment takes this blockage into account while requiring that positive directional stability is maintained at larger sideslip angles. The actual demonstration has been increased from a maximum range of ±10° at all speeds, as the previous amendment requires, to ±25° at slow speeds and linearly decreasing to ±10° at VNE. Alternatively to the previous range specified, the requirement limits the maximum sideslip to be demonstrated to at least 0.1g of sideforce or the maximum sideslip attained when full directional control is applied. As in the previous amendment, sufficient cues should alert the pilot when approaching sideslip limits.
b. Procedures. The policy material pertaining to the procedures outlined in this section remain in effect.
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AC 27-1B, Chg 4 3/31/2014 [This page intentionally blank.] Page B – 60.8
SUBPART B - FLIGHT
GROUND AND WATER HANDLING CHARACTERISTICS AC 27.231. § 27.231 GROUND AND WATER HANDLING CHARACTERISTICS-- GENERAL.
a. Explanation. The rule states: “The rotorcraft must have satisfactory ground and water handling characteristics, including freedom from uncontrollable tendencies in any condition expected in operation.” In addition, §§ 27.235, 27.239, and 27.241 contain specific requirements concerning ground and water handling characteristic evaluations.
b. Procedures.
(1) During the flight test program and the F&R program (§ 21.35(b)(2)), the rotorcraft will be subjected to evaluations at various weight and CG conditions. Any uncontrollable tendencies found during these test programs must be corrected.
(2) Controllable or damped vibrations or oscillations on the ground or in the water are acceptable, provided the design limits of the rotorcraft are not exceeded.
(3) Any significant vibration or oscillation characteristics found during tests should be described in the test report, and the rotorcraft flight manual should contain appropriate descriptions and procedures to describe and either avoid or handle significant characteristics.
(4) For rotorcraft equipped with wheel gear, the evaluation should include takeoff, landing, and taxi at the maximum speed and at CG extremes. If a nose or tail wheel lock/swivel control is installed, each position should be evaluated for limiting takeoff, landing, and taxi speeds. Maximum substantiated speed values should be included in the RFM as limitations.
(5) For water operations, the wave height and frequency or “sea state” should be included as a limitation or, if no limit was reached during testing, the demonstrated values should be placed in the Performance Section of the RFM. Information or limits on the allowable “sea state” for rotor startup and shutdown should also be included.
AC 27.235. § 27.235 TAXIING CONDITION .
a. Explanation. The rotorcraft is designed for certain landing load factors (§§ 27.471 and 27.473). The rotorcraft must not attain a load factor in excess of the design load factor when taxied over the roughest ground that may reasonably be expected in normal operation at the expected taxi speeds. This rule applies to wheel landing gear equipped rotorcraft.
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b. Procedures. The structural substantiation data contain the allowable design limits for the rotorcraft. A calibrated accelerometer or load factor “g” meter should be installed as near as practicable to the rotorcraft CG to record the maximum vertical load factor attained. Instrumentation of the landing gear and/or related structure may also be an acceptable means of showing compliance.
(1) Calibrated instrumentation should be installed to record the maximum loads or maximum vertical load factor attained during the taxi tests.
(2) The taxi surface should be evaluated for compliance with the rule.
Corrugated surfaces as well as broken or uneven surfaces (in accordance with the rule) should be used.
(3) Representative typical taxi speeds, up to the maximum selected by the applicant, should be attained over the selected taxi surfaces.
(4) A light and heavy rotorcraft weight condition should be evaluated.
(5) Limitations appropriate for the rotorcraft design should be included in the flight manual. If these tests indicate that it is unlikely that limit load factors will be attained while taxiing, flight manual limitations may not be necessary.
(6) Pertinent taxi information obtained from these test conditions may be included in normal procedures of the flight manual.
AC 27.239. § 27.239 SPRAY CHARACTERISTICS.
a. Explanation. The intent of this requirement is to evaluate by demonstration that water spray does not obscure visibility (day or night) or damage the rotorcraft during normal waterborne operation (for those rotorcraft which have waterborne or amphibious capability).
b. Procedures.
(1) The following maneuvers should be evaluated in ambient conditions up to the proposed sea state or wave height for operation.
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Con- Rotor Altifig. Condition Weight CG RPM tude Remarks 1 Taxi Max Optional Max SL Speeds up to maximum proposed for water operation.
2 Hover Max Opt Max - Determine critical hover height, if any.
3 Takeoff Max Opt Max SL Unstick at maximum proposed water operation speed.
4 Land Max Opt Max SL Touchdown at maximum proposed for water operation.
5 Shutdown Opt Opt - SL Shut down the rotorcraft.
6 Start Max Opt Max SL Start engines and release rotor brake.
(2) The maximum sea state or wave height evaluated under this rule should be stated and included in the limitations section of the flight manual.
(3) The effect of saltwater contamination and deterioration of turbine engines and other component parts of the rotorcraft should be considered in accordance with § 27.609 and paragraph AC 27.609. Information on saltwater effect and attendant corrective action should be provided in the flight manual, if appropriate, and in the maintenance manual.
AC 27.241. § 27.241 GROUND RESONANCE.
a. Explanation.
(1) The rule states: “The rotorcraft may have no dangerous tendency to oscillate on the ground with the rotor turning.” This rule is a flight requirement that pertains to demonstrating freedom from dangerous oscillations on the ground. CAR Part 6, predecessor to FAR Part 27, originally contained a “strength requirement” under § 6.203 requiring ground vibration tests. These tests would identify critical vibration frequencies and modes of the rotorcraft. CAR Part 6, Amendment 6-4, effective October 1, 1959, removed this ground vibration requirement because the agency concluded that if any major component has a natural frequency which could be excited by some operating parameter, such a condition would be revealed in the course of other ground and flight tests. The FAA/AUTHORITY apparently was depending on demonstrations under § 6.131/§ 27.241 and the flight load survey data (§ 27.571) to satisfy the objective of the vibration test. However, Part 27, Amendment 27-2, contained new § 27.663 adding reliability and damping action investigation Page B - 63
requirements for ground resonance prevention means. A ground vibration survey was not reinstituted by the adoption of § 27.663. Compliance with § 27.663 does require investigation and substantiation as stated.
(2) “Ground resonance” is a mechanical instability of the aircraft while in contact with the ground, often when partially airborne. Stated another way “ground resonance” is a self-excited mechanical instability that involves coupling between the in-plane motion of the rotor blade and the motion of the rotorcraft as a whole on its landing gear (reference “Aerodynamic of the Helicopter,” Gessow & Myers, page 308).
It is caused by the motion of the blade in the plane of rotation (called in-plane vibration) coupled with a rocking or vertical motion of the aircraft as a whole. The tires, landing gear, and rotor pylon restraint structure act as a spring with a vibration frequency which coincides or couples with the natural in-plane frequency of the blade about a real or effective drag hinge in the plane of rotation. When the frequencies of the two mot ions (rotor and airframe) approach each other and couple, a violent shaking of the rotorcraft may occur which, if undamped, could result in the destruction of the rotorcraft.
(3) Ground resonance can occur due to flexibility in the rotor pylon restraint system as well as with landing gear flexibilities. This mode of vibration or resonance can happen in flight (called air resonance) as well as on the ground and should be addressed in the certification program. The evaluation should include variations in stiffness and damping that could occur in service to the rotor pylon restraints.
(4) Ground resonance may be prevented by placing the first order in-plane vibration frequency above the rotor turning speed.
(5) For such configurations which are not susceptible to ground resonance (first order in-plane frequency above rotor turning speed), a simple rotor RPM run-up and run-down with appropriate cyclic control displacement (i.e., excitation of any inherent vibrations) is adequate demonstration that a ground resonance condition does not exist.
Unhinged “rigid” rotors, such as Bell Helicopter two-blade designs, are this type of rotor system.
(6) For configurations that are susceptible to ground resonance (i.e., first in-plane frequency is below the rotor turning speed), ground resonance is generally prevented by dampers on the blade acting in the plane of rotation, dampers on the landing gear (sometimes serving as oleo struts), or proper placement of the landing gear frequencies combined with rotor and/or landing gear dampers.
(7) Elastomeric components (in the rotor pylon support system, possibly in the landing gear, and possibly in the rotor head) are significantly affected by ambient temperature prior to warm-up. Their damping characteristics require thorough investigation for the range of rotorcraft operating environment as noted in § 27.663.
b. Procedures.
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(1) Under all conditions, any oscillations which may be introduced should be damped. However, no instability should occur at any operating condition such as during RPM changes from minimum to maximum and idle to maximum. For rotorcraft with wheel gear, uneven taxi surfaces in conjunction with particular taxi speeds, may excite ground resonance and should be evaluated by taxiing on typical surfaces. This evaluation may be conducted in conjunction with the tests of § 27.235. In operation, the resonance characteristics should be checked during takeoff and landing at zero speed and during run-on landings using various power values.
(2) For those aircraft equipped with Stability Augmentation Systems (SAS), all ground resonance investigations should be conducted with SAS on and SAS off. This includes the hovering and running takeoffs and landings, taxi tests, and specific ground resonance tests noted herein. Consideration should be given to conducting tests in various SAS configurations such as roll channel on and pitch channel off, where such configurations are possible and authorized.
(3) For each rotorcraft configuration tested, the aircraft should be positioned on the ground in flat pitch with the rotor stabilized at the minimum practical rotational speed or optionally at a speed shown analytically to have significant margin from indicated resonant conditions. Control system inputs should be used to disturb the system for evaluation of subsequent damping.
(4) For each incremental increase in rotor speed and for each rotor speed setting at increments of collective pitch settings, cyclic and collective inputs should be investigated prior to proceeding to the next rotor speed setting. These inputs should cover the appropriate range and combinations of amplitude and frequency. The collective pitch setting increments should range from flat pitch to light on the landing gear prior to fully airborne, depending upon the test sequence for minimum risk.
(5) Cyclic pitch inputs should be made either by the pilot through the cyclic stick or through a signal-generating device working in conjunction with the cyclic controls.
For each frequency of input, amplitude of the inputs should be increased incrementally and ultimately should be large enough to generate responses representative of normal ground and flight operation on the rotor and support system. The inputs should continue for a time sufficient to obtain representative responses, typically time sufficient to execute five complete circles of the cyclic stick (about neutral) at the selected frequency.
(6) The excitation frequency should be such as to excite the blade in-plane frequency. Rotor speed settings should be increased to 1.05 times the maximum power-on rotor speed. Collective pitch settings should be increased in increments of not more than 20 percent to maximum collective or alternately to the collective setting required to become partially airborne (when the cyclic is displaced as noted).
(7) Typically, articulated rotor aircraft have natural frequencies on the blade in lag of approximately 0.3 times the power-on main rotor RPM. Soft in- plane rotors have Page B - 65