(3) Emergency Procedures. Emergency procedures are concerned with foreseeable but unusual situations in which immediate and precise action by the crew, as detailed in the recommended procedures, may be expected to reduce substantially the risk of disaster. Typical examples of incidents considered to be emergencies are fire, ditching, loss of tail rotor thrust, etc.
(4) Ditching Procedures. Amendment 29-12 added ditching standards to Part 29. When ditching approval is requested, appropriate procedures and information will be included in the manual. Scale model tests are generally used to prove autorotation "ditching" characteristics and to prove stability in the water (capsize threshold) of the rotorcraft type design. Many rotorcraft designs require emergency float bags that deploy either before water contact or shortly after water contact to provide the flotation and stability necessary to comply with the requirements.
Autorotation altitudes and airspeeds and water contact information, if appropriate, derived from or used during the ditching model tests, should be confirmed during FAA/AUTHORITY flight tests and should be included in the manual. Information concerning sea states or wave heights to length ratios, investigated and found satisfactory, may be included in the manual if nonsevere sea states are likely to be exceeded.
(i¡) Instructions for deploying liferafts may be needed for certain designs.
For example, if liferafts are stowed outside the cabin, special instructions may be necessary.
(5) Evacuation Procedures for Rotorcraft Litter Configurations. Appropriate procedures and minimum crew requirements should be considered and included in the manual or manual supplement, if necessary, to assure timely evacuation.
(6) The use of illustrations to show controls, instruments, explain systems, etc., is encouraged.
765. § 29.1587 (Amendment 29-24) PERFORMANCE INFORMATION.
a. Explanation.
(1) This section should contain the performance information necessary for operation in compliance with applicable performance requirements of Part 29 and applicable special conditions, together with additional information and data essential for implementing pertinent operational requirements.
(2) Performance information and data may be presented for the range of weight, altitude, temperature, and other operational variables stated as operational performance limitations. Performance information which exceeds any operating limitation should be shown only as required for clarity of presentation. If data beyond 1156 Par 764
7/30/97 АС 29-2B operating limits are shown, the limits should be clearly marked and the data outside of the limits clearly distinguishable from the data within the limits.
(3) Performance information presented in the unapproved or manufacturers' data" section of the RFM should not include performance data that are beyond operating limitations unless the particular operating limit that may be exceeded is clearly distinguishable from similar performance data that are within limits. For example, if the weight-altitude-temperature (WAT) limits for takeoff and landing are based on in-ground-effect (IGE) hover performance capability at a 5-foot skid height, 3-foot skid height hover performance data allowing increased hovering weights should not be presented in the manufacturers' data unless clearly identified as being beyond operating limitations for normal operations. It is recommended that performance information and data be presented substantially in accordance with the following paragraphs. Where applicable, reference to the appropriate requirement of the certification or operating regulation should be included.
General. Include all descriptive information necessary to identify the configuration and conditions for which the performance data are applicable. Such information may include the complete model designations of rotorcraft and engines, definition of installed rotorcraft features, and equipment that affects performance together with the operative status thereof. This section should also include definitions or terms used in the performance section (i.e., IAS, CAS, ISA, configuration, CDP, Vross, Category A, Category B, LDP, etc.) plus calibration data for airspeed, altimeter, ambient air temperature, and other information of a general nature.
(i¡) Performance Procedures. The procedures, techniques, and other conditions associated with obtainment of the flight manual performance should be included. The procedures may be presented as a performance subsection or in connection with a particular pertormance graph. In the latter case, a comprehensive listing of the conditions associated with the particular performance may serve the objective of "procedures" if sufficiently complete. Performance figures are based on the installed minimum specification engine, unless normally depreciated engine performance is approved.
(iii) Wind Accountability. Wind accountability may be utilized for determining takeoff and landing field lengths. This accountability may be up to 100 percent of the minimum wind component along the takeoff or landing path opposite to the direction of takeoff. Wind accountability data presented in the RFM should be labeled "UNFACTORED" (if 100 percent accountability is taken) and should be accompanied by the following note: "Unless otherwise authorized by operating regulations, the pilot is not authorized to credit more than 50 percent of the performance increase resulting from the actual headwind component and must reduce performance by 150 percent of the performance decrement resulting from the actual tail wind component." In some rotorcraft, it may be necessary to discount the beneficial aid to takeoff performance for winds from zero to 10 knots. This should be done if it is Par 765 1157
AC 29-2B 7/30/97/ evident that the winds trom zero to 10 knots have resulted in a significant degradation to the takeoff performance due to flight through the main rotor vortex. Degradation may be determined by determining the power required to fly, by reference to a pace vehicle, at speeds of 10 knots or less.
(iv) The following list is illustrative of the information that should be provided for a transport Category "A" and "B" rotorcraft.
(A) Density altitude chart for converting from pressure to density altitude.
(B) Temperature conversion chart (°C to °F to °C).
(C) Airspeed calibration (calibrated vs. indicated airspeed) for both pilot and copilot systems for level flight, climb, autorotation, and recommended approach rate of descent.
(D) Altimeter correction for pilot and copilot instruments showing the correction factor vs. indicated airspeed at sea level and altitude.
(E) Hover performance charts both in and out-of-ground (OGE) effect with instructions for their use. The OGE hover performance chart is not required but may be useful.
(F) A series of climb performance charts for various weights showing rate of climb vs. pressure altitude for a range of temperatures and showing the variation of best rate of climb speed with pressure altitude. The conditions should appear on each chart (i.e., power, weight, single, or multiengine, etc.). The OEI climb performance charts at 30-minute power and maximum continuous power or at continuous OEI power should provide rate of climb performance down to a minimum of -500 feet/min. The effect of engine air bleed, particle separators or other devices, on the rate of climb/descent performance must be provided.
(G) A chart showing the takeoff flight path for Category A presented in height vs. distance from the hover wheel height to the point at which Voss and not less than 35 feet is reached, and the rejected takeoff distance. The chart should identify the critical decision point and Voss.
(H) Charts to allow calculation of distance to climb at Voss from the point at which Vross and not less than 35 feet is reached (or from the lowest point of the takeoff profile for elevated heliport) to 200 feet with one engine inoperative and other engines within approved operating limitations. If conservative, providing charts to allow calculation of the total distance from Vross and 35 feet to Vy and 200 feet is allowed.
1158 Par 765
(I) A series of charts to allow calculation of any additional distance which may be required to accelerate to best rate of climb speed from Vross with one engine inoperative and other engines within approved operating limitations. If conservative, providing charts to allow calculation of the total distance from Vross and 35 feet to Vy and 200 feet is allowed.
(J) Charts to allow calculation of distance to climb at Vy from 200 feet to 1000 feet above the takeoff surface (or from the lowest point of the takeoff profile for elevated heliport) with one engine inoperative and other engines at 30-minutes OEI power or maximum continuous OEI power. If conservative, providing charts to allow calculation of the total distance from Vross and 35 feet to Vy and 1000 feet is allowed.
(K) Landing distance chart for Category A showing the landing distance from a 50-foot height (25-foot for VTOL operations from an elevated heliport) to a stop with one engine inoperative vs. pressure altitude over the range of temperatures being certified. This chart should identify the balked landing decision point (LDP) so the pilot will know how to achieve this performance.
(L) For Category B, a series of charts at various weights showing takeoff distance from hover to 50 feet vs. pressure altitude over the range of temperatures being certified.
(M) For Category B, a landing distance chart similar to the one for Category A from a 50-foot height to stop with one engine inoperative.
(N) For turbine-powered rotorcraft in all categories, a power assurance check chart.
(O) For Category B, a statement of the maximum crosswind and downwind components that have been demonstrated as safe for operation near the ground unless this information is incorporated as an operating limitation. (See Paragraph 763 of this AC.)
(P) For Category B, the height-velocity (HV) envelope except for rotorcraft which must incorporate the HV diagram as an operating limitation.
(Q) For Category B, the autorotative glide distance as a function of altitude if required by § 29.71. (See Paragraph 68 of this AC.)
(V) Miscellaneous Performance Data. Any performance information or data not covered in items (A) through (Q) above, but considered necessary to enhance safety or to enable application of the operating regulations, should be included.
Par 765 1159
765A. § 29.1587 (Amendment 29-40) PERFORMANCE INFORMATION.
a. Explanation. Amendment 29-40 added a requirement to provide the steady gradient of climb for each weight, altitude, and temperature for which Category A performance is presented. No minimum climb gradient has been required.
b. Procedures. No additional flight testing is required beyond that for compliance with the Category A performance requirements. Climb gradient data should be calculated and presented for all weights, altitudes, and temperatures for which takeoff data is scheduled. Gradients should be established for the first and second segment climb under the conditions specified in § 29.67(a)(1) and (a)(2).
766. § 29.1589 LOADING INFORMATION.
a. Explanation. Control of the rotorcraft weight and balance is an operational function, and is the responsibility of the operator. However, instructions necessary to enable loading of the rotorcraft within the established limits of weight and center of gravity, and to maintain the loading within such limits are required by the operating regulations, and inclusion of such loading instructions in the Rotorcraft Flight Manual is required by § 29.1583(c). Approved loading instructions, therefore, must be presented in the Rotorcraft Flight Manual, and at the option of the applicant, may be included in the approved portion or may be included in the unapproved portion.
b. Procedures.
(1) For the purpose of the flight manual, distinction is made here between the loading instructions required by the certification requirements of Part 29, and the weight and balance data required by the operating requirements. The former prescribed information is applicable to the rotorcraft type, and is subject to FAA/AUTHORITY approval as flight manual material.
(2) For compliance with the noted requirements, it is necessary for the applicant to develop weight and balance data and loading instructions as necessary to satisfy the needs of both certification and operation. In order to consolidate in one document information on rotorcraft loading, however, it is recommended that the weight and balance data be developed to include appropriate loading instructions, and that both be included in the Rotorcraft Flight Manual as an "unapproved" section entitled, "Weight and Balance." Such a section should include the following statement as a note: "In accordance with FAA/AUTHORITY procedures, the detail weight and balance data of this section are not subject to FAA/AUTHORITY approval. The loading instructions of this section, however, have been approved by FAA/AUTHORITY as satisfying all requirements for instructions on loading of the rotorcraft within approved limits of weight and center of gravity, and on maintaining the loading within such limits."
1160 Par 765A
(3) An actual or specimen weight and balance section should be included in the initial submittal of the manual. Weight and balance data for each particular rotorcraft need not be submitted as flight manual material.
(4) The weight and balance material outlined below is believed to be adequate for rotorcraft with conventional loading and fuel-management techniques. For rotorcraft which necessitate redistribution of fuel (other than normal consumption) to maintain loading within prescribed limits, the material should be amplified as necessary.
Weight Limits. Contained in limitations section of the flight manual.
(i) Center of Gravity Limits. Contained in the limitations section of the flight manual.
(iii) Dimensions and Datum Line Locations. The dimensions and relative location of rotorcraft features associated with weighing and loading of the rotorcraft and with weight and balance computations should be described and/or illustrated.
(iv) Equipment List. The rotorcraft should be defined or described sufficiently to identify the presence or absence of optional systems, features, or installations that are not readily apparent. In addition, all other items of fixed and removable equipment included in the empty weight should be listed.
(v) Fuel and Other Liquids. Fuel and other liquids, including passengerservice liquids that are included in the empty weight, should be identified and listed together with information necessary to enable ready duplication of the particular condition.
(vi) Weight Computations. Computations of the empty weight and emptyweight CG location should be included.
(vii) Empty Weight and Empty-Weight Center of Gravity Location.
Statement of these values should be included.
(viii) Loading Schedule. Loading schedule should be included, if appropriate.
(ix) Loading Instructions. Complete instructions relative to the loading procedure, or to use the loading schedule, must be included.
(x) Special Consideration. Consideration should be given to the lateral center-of-gravity loading instructions when various kits such as a side mounted hoist are installed.
Par 766 1161
767.-774. BESERVED.
Intentionally Left Blank 1162 (thru 1170) Par 767
CHAPTER 3
MISCELLANEOUS AIRWORTHINESS STANDARDS
775. JER CERTIFIÇATION
a. Explanation. Requirements for instrument flight rules (IFR) have been incorporated into Part 29, Appendix B, utilizing a regulatory format. Various information from previous interim standards, procedures, test techniques, and acceptable means of compliance for rotorcraft IF flight are included in the following sections.
b. Procedures.
(1) General.
The certified instrument flight envelope may be more restrictive than the visual flight rules (VFR) envelope in terms of weight, center of gravity, speed, altitude, or rate of climb and descent. The approved envelope should be operationally practical such that it does not impose constraints with which the crew has difficulty complying.
Controllability requirements are to be met from 0.9 VMINI to 1.1 VNE:
Stability requirements must be met where specified. Stability devices are to be designed to allow safe flight following a failure. The evaluating pilot should assure that all equipment and devices installed for IFR, including reasonable failures of that equipment, do not compromise the VFR approval for that rotorcraft. An example of this would be a stability system failure that caused loss of swashplate or tail rotor control travel when failed in a hardover condition. If the device remains in the hardover position after the stability system is turned off, control capability may be compromised.
Cyclic controllability tests at high speed and at the limiting rearward flight condition, or tail rotor tests in sideward flight at high altitude, may reveal a lower control capability and a more restrictive envelope. In addition, controllability testing should be accomplished with the control rigging set at the most adverse production tolerance for the test condition; e.g., minimum forward swashplate for high speed testing.
(2) Trim. Compliance with the IF trim requirement may be met by use of a magnetic brake with a recentering button, an electrically driven trim system activated by a "beeper" type control, or other means, so long as the system does not introduce any objectionable discontinuities in the force gradient or otherwise result in objectionable flight characteristics. Trim release devices should be free of objectionable stick jump.
Electrically driven trim systems should have a smooth change in force with a rate compatible with the normal rotorcraft maneuvers. Only the cyclic trim control must exhibit positive self-centering characteristics. Collective and directional controls are not Par 775 1171
required to incorporate positive self-centering characteristics, but these controls should not move when released by the pilot (adjustable friction devices are satisfactory);
however, for systems which use hydraulic or pneumatic dampers, control motion following release by the pilot is permitted during the time interval when the damper is bleeding off. Movement of the trim controls should produce a similar effect on the rotorcraft in a plane parallel to that of the control motion. The control system free play and breakout force must be evaluated to assure a close and direct correlation between control input (force and deflection) and rotorcraft response (pitch, roll, yaw, and heave (vertical motion)), and to permit small, precise changes in flight path. If trim control is provided in a stability augmentation system (SAS), the control should be of such design and so installed that any failure will not create a hazardous condition. If an inadvertent out-of-trim condition can be developed, its effect on the rotorcraft should be investigated. These failures or malfunctions should be investigated as outlined in
(6) "Stability Augmentation Systems" which follows. The controls for this trim function should be installed such that, the controls should operate in the plane and with the sense of motion of the rotorcraft. Each control means should have the direction of motion plainly marked thereon or adjacent to the control.
(3) Static Longitudinal Stability.
Positive static longitudinal stability is a key IFR requirement which assures a self-correcting airspeed response and allows a pilot to recognize any substantial change in speed. The phrase "substantial speed change" as used in FAR 29, Appendix B, Paragraph IV, is normally considered to mean at least a 10 knot departure from trim speed. Such a change in airspeed must be accompanied by a stick force clearly perceptible to the pilot (l.e., a discernible and quantifiable force gradient).
Very shallow force gradients can be approved for systems with low deadband and low friction. Systems with significant friction and deadband require much steeper force gradients to be acceptable. The longitudinal force gradient can be determined by either of two methods. The most commonly used method (applicable only to irreversible control systems) measures the cyclic forces with the rotorcraft on the ground and the rotor stopped (with hydraulic and electric power units if required). The force applied to the cyclic stick and the cyclic stick displacement are measured and a plot of stick force versus displacement in each longitudinal direction is obtained. Following the ground test, the longitudinal static stability tests are conducted in flight as described in Paragraph 85. The cyclic displacement measurements gathered during flight test are then assigned force values from the ground mechanical characteristics test and the force values are cross plotted with the corresponding airspeeds to produce a plot of cyclic force versus airspeed. The trim system should be on during the test and the aircraft trimmed at the trim speed. After each end point, the cyclic should be allowed to slowly return to the trim position. When all the force is released from the cyclic stick and the airspeed has stabilized, note the airspeed. An alternate method of determining the longitudinal stick force stability is to measure the force on the cyclic stick in flight using a hand held force gage or other force measuring instrumentation. The in-flight technique is the same as the first method. Testing should be accomplished at a 1172 Par 775
minimum of two altitudes. One altitude should be low enough to assure limiting power is attained. Another should be at or near the maximum approved altitude. Reasonable interpolation is allowed. If no marginal areas are apparent interpolation over a 10,000-foot altitude range is considered reasonable.
(ii) Tests for static longitudinal stability during approach should include the steepest approach gradient for which approval is requested. Static stability tests may be simulated by initially establishing a trimmed rate of descent for maximum approach gradient assuming zero wind conditions. Actual approach tests at the maximum approved gradient should be conducted to evaluate tracking and maneuverability, including the capability to correct downward to a glide path when approaching in a slight (10 knot) tailwind condition.
(4) Static Lateral - Directional Stability.
(i) Tests for directional stability usually require instrumentation for lateral cyclic position, pedal position, and sideslip angle. Testing for compliance with the specific directional requirement is relatively simple; however, the pilot should look tor significant longitudinal trim changes and short period dynamic modes which might occur only during sideslip conditions. Side force characteristics are indicated by the variation of bank angle with sideslip during steady heading sideslips. The number of ball widths of deflection is also indicative of the side force cue available to the pilot. A correlation between sideslip angle and ball widths of skid can be obtained at given speeds for use during later testing after sideslip instrumentation is removed. A simple yaw string can be calibrated in a similar manner. The TIA should define the maximum sideslip angles which should not be exceeded during the flight test program. These angles must not be greater than the structural sideslip envelope substantiated and are not required to be that sideslip angle obtained with full directional pedal deflection. Sufficient side force cues should accompany sideslip to alert the crew when approaching sideslip limits.
This is needed to assure that structural sideslip limits will not be inadvertently exceeded in service. Although not stated in the requirement, flight conditions for demonstration of static longitudinal stability are also appropriate for demonstration of static lateraldirectional stability.
(®) Dihedral requirements may be more difficult to assess. For those rotorcraft which do not meet the position and force gradient requirements for the conventional, cross-controlled sideslips, there are alternative tests which may be used to determine acceptable characteristics. If directional pedals are utilized in steady sideslips, the resultant rolling tendency is the sum of (1) the aircraft's roll due to sideslip tendency (dihedral) and (2) the aircraft's roll due to directional control input. If the rotorcraft has a tail rotor which is excessively high or low in relation to the rotorcraft's vertical center of gravity (CG), application of tail rotor thrust will introduce a significant rolling moment. The basic intent of dihedral stability testing is to determine the rotorcraft response to sideslip exclusive of directional control input. In general, if a tail rotor configuration is involved, and the tail rotor is above the vertical CG of the Par 775 1173
АС 29-2B 7/30/97 rotorcraft, the effect of pedal input upon dihedral effect is destabilizing during conventional, control-induced sideslips.
(™1) There are two alternate methods which, for small angles of sideslip, can give an indication of the basic dihedral stability of the rotorcraft. Both methods involve freezing directional controls while artificially creating sideslip by other means.
(iv) The first method is only applicable for rotorcraft with single main rotor systems. To utilize this method, the rotorcraft is stabilized in a given flight condition and small collective (torque) changes are applied in each direction (e.g., +5% & 110%) while holding pedals fixed. Sideslip angle, lateral control position, and lateral control force may be measured and plotted for small torque changes from trim. This technique will not work for aircraft which have collective to pedal or collective to lateral control couplings.
(V) In the second method, the rotorcraft is stabilized in a trimmed flight condition with a small amount of bank (5-10°). The rotorcraft is then rolled to an approximately equal angle of bank in the opposite direction holding the pedals fixed.
The change in direction of bank results in a small change in sideslip angle and again sideslip angle may be plotted versus lateral control position and/or force. This test should be conducted in both directions and the results averaged. This method can give reasonably accurate results for small perturbations. Other factors contribute to the results of either of these two methods. It is always important to assess the roll due to sideslip tendency with pedal induced sideslips to assure lateral control forces are reasonable and in a proper direction for directional out-of-trim conditions, and to assure the pilot has adequate sideslip cues.
(vi) Wording of the dihedral requirement is intended to allow slightly negative dihedral stability at critical loading conditions. This will ordinarily result in positive dihedral stability throughout a great majority of the approved loading envelope.
The test for maximum allowable negative dihedral effect would involve stabilization at a required flight condition, inducing a sideslip up to t10° from trim, then assessing lateral cyclic friction/deadband to determine if roll is restrained while remaining in the control system friction/deadband so that the control may be released without resulting in the aircraft rolling in the adverse direction. When testing for this condition, lateral cyclic friction should be adjusted to the minimum value.
(vii) The intent of the dihedral rule is to allow small amounts of control system friction and deadband to mask small values of negative dihedral. Where slope of the negative dihedral versus sideslip exceeds these small values, the negative dihedral shall not be approved. The operational pilot must not be presented with opposite cyclic sensing for similar sideslip conditions as loadings and flight conditions change. In general, large values of control system friction and deadband are 1174 Par 775
undesirable. The addition of friction or deadband into the control system for the purpose of satisfying the dihedral requirement is not acceptable.
(vill) In approving small negative dihedral values, the pilot should ensure that other positive flight cues, such as suitable side force, accompany sideslip. This will aid the pilot in determining direction of sideslip so that no reverse sensing or confusion accompanies sideslip conditions.
(5) Dynamic Stability.
Dynamic characteristics are defined in quantitative terms; however, some areas of interpretation and technique need special consideration:
(A) Unlike fixed-wing aircraft where the size of the input has no effect on damping ratio, rotorcraft can be sensitive to the type and size of input used to excite each dynamic mode. For instance, it has been found that for the phugoid-type dynamic oscillation, damping ratio is inversely proportional to the size of the input. It therefore becomes important that dynamic excitations be sized to approximate the response of the rotorcraft in a moderate turbulent gust. Also, the dynamic input should be made with the control(s) which most accurately simulates the typical aircraft gust response. Obviously, for this evaluation some flying of the rotorcraft in turbulence is necessary to obtain knowledge of the rotorcraft's gust response. Pulses and doublets may be used to generate disturbances similar to a gust. To assist returning the controls) to the trim position a hand held jig may be used. Use of attitude and rate instrumentation is desirable. The pilot may find that collective excitation, or collective in conjunction with cyclic, is most appropriate for gust simulation.
(B) The second area of concern in evaluating dynamic response is whether to let only one axis respond to an excitation or to let the rotorcraft respond in two or more axes. When it can be done sately, the rotorcraft should be allowed to follow its dynamic response in all axes. In other words, if pitch oscillations feed into roll, the pilot should attempt to observe and record the total aircraft dynamic response in both pitch and roll.
(C) The third area concerns strict compliance with the exact wording of the dynamic requirement. In this regard, a neutrally damped oscillation with a period of 19 seconds would not be acceptable; however, a very divergent oscillation that doubles in amplitude in 21 seconds would be acceptable. The 19-second oscillation is much less severe than the 21-second oscillation and yet is unacceptable by the "letter of the law." Figure 775-1 below is a graphical display of the dynamic requirement. The 19- and 21-second oscillations are shown as points (1) and (2). Point No. 1 is positioned much more toward the acceptable portion of the graph and yet by the "letter of the lawi is unacceptable. The intent of the dynamic requirement is roughly approximated by the dashed-curved line. Areas to the right of that line may be considered for findings of equivalent safety.
Par 775 1175