AC 25.629-1C
7.1.2.2 The flutter analysis methods most extensively used involve the modal analysis with unsteady aerodynamic forces derived from various two- and three-dimensional theories. These methods are generally for linear systems. Analyses involving control system characteristics should include equations describing system control laws in addition to the equations describing the structural modes.
7.1.2.3 Airplane lifting surface divergence analyses should include all appropriate rigid body mode degrees-of-freedom since divergence may occur for a structural mode or the short period mode.
7.1.2.4 Loss of control effectiveness (control reversal) due to the effects of elastic deformations should be investigated. Analyses should include the inertial, elastic, and aerodynamic forces resulting from a control surface deflection.
7.1.3 Damping Requirements.
7.1.3.1 There is no intent in this AC to define a flight test level of acceptable minimum damping.
7.1.3.2 Flutter analyses results are usually presented graphically in the form of frequency versus velocity (V-f, figure 3) and damping versus velocity (V-g, figure 4 and figure 5) curves for each root of the flutter solution.
Figure 3. Frequency versus Velocity Frequency Velocity VD 1.15 VD
(5)
(1)
(4)
(3)
(2)
7.1.3.3 Figure 4 details one common method for showing compliance with the requirement for a proper margin of damping. It assumes that the structural damping available is 0.03 (1.5 percent critical viscous damping) and is the same for all modes as depicted by the V-g curves shown in figure 4. No 11
AC 25.629-1C
significant mode, such as curves (2) or (4), should cross the g=0 line below VD or the g=0.03 line below 1.15 VD. An exception may be a mode exhibiting damping characteristics similar to curve (1) in figure 4, which is not critical for flutter. A divergence mode, as illustrated by curve (3) where the frequency approaches zero, should have a divergence velocity not less than 1.15 VD.
Figure 4. Damping versus Velocity—Method 1 VD Velocity
(1)
(3)
(2)
(5) Damping g
0.03
(4) 0
1.15 VD
7.1.3.4 Figure 5 shows another common method of presenting the flutter analysis results and defining the structural damping requirements. An appropriate amount of structural damping for each mode is entered into the analysis prior to the flutter solution. The amount of structural damping used should be supported by measurements taken during full scale tests. This results in modes offset from the g=0 line at zero airspeed and, in some cases, flutter solutions different from those obtained with no structural damping. The similarity in the curves of figure 4 and figure 5 are only for simplifying this example. The minimum acceptable damping line applied to the analytical results as shown in figure 5 corresponds to 0.03 or the modal damping available at zero airspeed for the particular mode of interest, whichever is less, but in no case less than 0.02. No significant mode should cross this line below VD or the g=0 line below 1.15 VD.
AC 25.629-1C
Figure 5. Damping versus Velocity—Method 2 VD Minimum Acceptable Damping Line Velocity
(1)
(3)
(2)
(5) Damping g 0
- 0.03
(4)
1.15 VD
7.1.3.5 For analysis of failures, malfunctions or adverse conditions being investigated, the minimum acceptable damping level obtained analytically would be determined by use of either method above, but with a substitution of VC for VD and the fail-safe envelope speed at the analysis altitude as determined by paragraph 6.1.3 of this AC.
7.1.4 Analysis Considerations.
Airframe aeroelastic stability analyses may be used to verify the design with respect to the structural stiffness, mass, fuel (including in-flight fuel management), automatic flight control system characteristics, altitude, and Mach number variations within the design flight envelope. The complete airplane should be considered as composed of lifting surfaces and bodies, including all primary control surfaces that can interact with the lifting surfaces to affect flutter stability. Control surface flutter can occur in any speed regime and has historically been the most common form of flutter. Lifting surface flutter is more likely to occur at high dynamic pressure and at high subsonic and transonic Mach numbers. Analyses are necessary to establish the mass balance and/or stiffness and redundancy requirements for the control surfaces and supporting structure and to determine the basic surface flutter trends. The analyses may be used to determine the sensitivity of the nominal airplane design to aerodynamic, mass, and stiffness variations. Sources of stiffness variation may include the effects of skin buckling at limit load factor, air entrapment in hydraulic actuators, expected levels of in-service freeplay, and control system components that may include elements with nonlinear stiffness. Mass variations include the effects of fuel density and distribution, control surface repairs and painting, and water and ice accumulation.
7.1.4.1 Control Surfaces.
Control surface aeroelastic stability analyses should include control surface rotation, tab rotation (if applicable), significant modes of the airplane, control surface torsional degrees-of-freedom, and control surface 13
AC 25.629-1C
bending (if applicable). Analyses of airplanes with tabs should include tab rotation that is both independent and related to the parent control surface.
Control surface rotation frequencies should be varied about nominal values as appropriate for the condition. The control surfaces should be analyzed as completely free in rotation unless it can be shown that this condition is extremely improbable. All conditions between stick-free and stick-fixed should be investigated. Freeplay effects should be incorporated to account for any influence of in-service wear on flutter margins. The aerodynamic coefficients of the control surface and tab used in the aeroelastic stability analysis should be adjusted to match experimental values at zero frequency. Once the analysis has been conducted with the nominal, experimentally adjusted values of hinge moment coefficients, the analysis should be conducted with parametric variations of these coefficients and other parameters subject to variability. If aeroelastic stability margins are found to be sensitive to these parameters, then additional verification in the form of model or flight tests may be required.
7.1.4.2 Mass Balance.
7.1.4.2.1 Applicants may evaluate the magnitude and spanwise location of control surface balance weights by analysis and/or wind tunnel flutter model tests.
If the control surface torsional degrees of freedom are not included in the analysis, then adequate separation needs to be maintained between the frequency of the control surface first torsion mode and the flutter mode.
7.1.4.2.2 Control surface unbalance tolerances should be specified to provide for repair and painting. The accumulation of water, ice, and/or dirt in or near the trailing edge of a control surface should be avoided. Freeplay between the balance weight, the support arm, and the control surface should not be allowed. Control surface mass properties (weight and static unbalance) should be confirmed by measurement before conducting a GVT.
7.1.4.2.3 The balance weights and their supporting structure should be substantiated for the extreme load factors expected throughout the design flight envelope. In the absence of a rational investigation, the following limit accelerations, applied through the balance weight center of gravity, should be used:
• 100g normal to the plane of the surface.
• 30g parallel to the hinge line.
• 30g parallel to the plane of the surface and perpendicular to the hinge line.
7.1.4.3 Passive Flutter Dampers.
Control surface passive flutter dampers may be used to prevent flutter in the event of failure of some element of the control surface actuation 14
AC 25.629-1C
system or to prevent control surface buzz. Flutter analyses and/or flutter model wind tunnel tests may be used to verify adequate damping. Damper support structure flexibility should be included in the determination of adequacy of damping at the flutter frequencies. Any single damper failure should be considered. Combinations of multiple damper failures should be examined when not shown to be extremely improbable. The combined freeplay of the damper and supporting elements between the control surface and fixed surfaces should be considered. Provisions for in-service checks of damper integrity should be considered. See paragraph 6.3.3 of this AC for conditions to consider where a control surface actuator is switched to the role of an active or passive damping element of the flight control system.
7.1.4.4 Intersecting Lifting Surfaces.
Intersecting lifting surface aeroelastic stability characteristics are more difficult to predict accurately than the characteristics of planar surfaces such as wings. This is due to difficulties both in correctly predicting vibration modal characteristics and in assessing those aerodynamic effects that may be of second order importance on planar surfaces but are significant for intersecting surfaces. Proper representation of modal deflections and unsteady aerodynamic coupling terms between surfaces is essential in assessing the aeroelastic stability characteristics. The in-plane forces and motions of one or the other of the intersecting surfaces may have a strong effect on aeroelastic stability; therefore, the analysis should include the effects of steady flight forces and elastic deformations on the in-plane effects.
7.1.4.5 Ice Accumulation.
Aeroelastic stability analyses should use the mass distributions derived from ice accumulation up to and including those that can accrete in the applicable icing conditions in 14 CFR part 25, Appendices C and O. This includes any accretions that could develop on control surfaces. The analyses need not consider the aerodynamic effects of ice shapes. For airplanes proposed to be approved for operation in icing conditions, all the part 25, Appendix C icing conditions and the Appendix O icing conditions for which certification is sought are applicable. For airplanes excluded from § 25.1420, no evaluation of appendix O icing conditions is required.
For airplanes not approved for operation in icing conditions, all the icing conditions in Appendices C and O are applicable since the inadvertent encounter discussed in paragraph 6.2.2.3 of this AC can occur in any icing condition. For all airplanes, the ice accumulation determination should take into account the ability to detect the ice and, if appropriate, the time required to leave the icing condition.
AC 25.629-1C
7.1.4.6 Whirl Flutter.
7.1.4.6.1 The applicant’s evaluation of the aeroelastic stability should include investigations of any significant elastic, inertial, and aerodynamic forces, including those associated with rotations and displacements in the plane of any turbofan or propeller, including propeller or fan blade aerodynamics, powerplant flexibilities, powerplant mounting characteristics, and gyroscopic coupling.
7.1.4.6.2 Failure conditions are usually significant for whirl instabilities. Engine mount, engine gear box support, or shaft failures that result in a node line shift for propeller hub pitching or yawing motion are especially significant.
7.1.4.6.3 A wind tunnel test with a component flutter model, representing the engine/propeller system and its support system along with correlative vibration and flutter analyses of the flutter model, may be used to demonstrate adequate stability of the nominal design and failed conditions.
7.1.4.7 Automatic Control Systems.
The interactions of the airplane’s automatic control systems coupling with the structural modes must be controlled to prevent the occurrence of any aeroelastic instability (§ 25.629), which includes prevention of aeroservoelastic instability. These control systems could include flight control systems, autopilots, yaw damper systems, modal suppression systems, or any other feedback system that could interact with the airplane’s structural modes. Aeroelastic stability analyses of the basic configuration should include simulation of any control system for which interaction may exist between the sensing elements and the structural modes. Where structural/control system feedback is a potential problem, the effects of servo-actuator characteristics and the effects of local deformation of the servo mount on the feedback sensor output should be included in the analysis. The effect of control system failures on the airplane aeroelastic stability characteristics should be investigated.
Failures that significantly affect the system gain and/or phase and are not shown to be extremely improbable should be analyzed to ensure stability.
For nominal conditions, the structural modes should have the analytical stability margins specified below for any single control system feedback loop at speeds up to VD/MD and may linearly decrease to zero margins (stable at nominal gain and phase) at the envelope described in § 25.629(b)(1). For control laws with limited operating envelopes, the control law design speed may be used in place of VD/MD, and these margins may linearly decrease to zero (stable at nominal gain and phase) at the design speed plus 25 ft/sec equivalent airspeed increment, expanded by 15% in equivalent airspeed at both constant Mach number and constant 16
AC 25.629-1C
altitude. If these margins are not used, then a technical justification should be provided for the use and acceptance of alternative criteria.
7.1.4.7.1 A gain margin of at least 6 dB and, separately,
7.1.4.7.2 A phase margin of at least ±60°.
7.2 Testing.
The aeroelastic stability certification test program may consist of ground tests, flutter model tests, and flight flutter tests. Ground tests may be used for assessment of component stiffness and for determining the vibration modal characteristics of airplane components and the complete airframe. Flutter model testing may be used to establish flutter trends and validate aeroelastic stability boundaries in areas where unsteady aerodynamic calculations require confirmation. Full-scale flight flutter testing provides final verification of aeroelastic stability. The results of any of these tests may be used to provide substantiation data, to verify and improve analytical modeling procedures and data, and to identify potential or previously undefined problem areas.
7.2.1 Structural Component Tests.
Stiffness tests or GVTs of structural components are desirable to confirm analytically predicted characteristics and are necessary where stiffness calculations cannot accurately predict these characteristics. Components should be mounted so that the mounting characteristics are well defined or readily measurable.
7.2.2 Control System Component Tests.
When reliance is placed on stiffness or damping to prevent aeroelastic instability, the following control system tests should be conducted. If the applicant performs the tests off the airplane, the test fixtures should reflect local attachment flexibility.
7.2.2.1 Actuators for primary flight control surfaces and flutter dampers should be tested with their supporting structure. These tests are to determine the actuator/support structure stiffness for nominal design and failure conditions considered in the fail-safe analysis.
7.2.2.2 Flutter damper tests should be conducted to verify the impedance of damper and support structure. Satisfactory installed damper effectiveness at the potential flutter frequencies should be assured. The results of these tests can be used to determine a suitable, in-service maintenance schedule and replacement life of the damper. The effects of allowable in-service freeplay should be measured.
7.2.3 Ground Vibration Tests.
7.2.3.1 Ground vibration tests or modal response tests are normally conducted on the complete conforming airplane. A GVT may be used to check the mathematical structural model. Alternatively, the use of measured modal data alone in aeroelastic stability analyses, instead of analytical modal data 17
AC 25.629-1C
modified to match test data, may be acceptable provided that the accuracy and completeness of the measured modal data is established. Whenever structural modifications or inertia changes are made to a previously certified design or a GVT-validated model of the basic airplane, a GVT may not be necessary if these changes are shown not to affect the aeroelastic stability characteristics.
7.2.3.2 The airplane is best supported such that the suspended airplane rigid body modes are effectively uncoupled from the elastic modes of the airplane.
Alternatively, a suspension method may be used that couples with the elastic airplane, provided that the suspension can be analytically de-coupled from the airplane structure in the vibration analysis. The former suspension criterion is preferred for all GVTs and is necessary in the absence of vibration analysis.
7.2.3.3 The excitation method needs to have sufficient force output and frequency range to adequately excite all significant resonant modes. The effective mass and stiffness of the exciter and attachment hardware should not distort modal response. More than one exciter or exciter location may be necessary to ensure that all significant modes are identified. Multiple exciter input may be necessary on structures with significant internal damping to avoid low response levels and phase shifts at points on the structure distant from the point of excitation. Excitation may be sinusoidal, random, pseudo-random, transient, or other short duration, non-stationary means. For small surfaces, the effect of test sensor mass on response frequency should be taken into consideration when analyzing the test results.
7.2.3.4 The minimum modal response measurement should consist of acceleration (or velocity) measurements and relative phasing at a sufficient number of points on the airplane structure to describe accurately the response or mode shapes of all significant structural modes. In addition, the structural damping of each mode should be determined.
7.2.4 Flutter Model Tests.
7.2.4.1 Dynamically similar flutter models may be tested in the wind tunnel to augment the flutter analysis. Flutter model testing can substantiate the flutter margins directly or indirectly by validating analysis data or methods. Some aspects of flutter analysis may require more extensive validation than others, for example, control surface aerodynamics, T-tails, and other configurations with aerodynamic interaction and compressibility effects. Flutter testing may additionally be useful to test configurations that are impractical to verify in flight test, such as fail-safe conditions or extensive store configurations. In any such testing, the mounting of the model and the associated analysis should be appropriate and consistent with the study being performed.
AC 25.629-1C
7.2.4.2 Direct substantiation of the flutter margin (clearance testing) implies a high degree of dynamic similitude. Such a test may be used to augment an analysis and show a configuration flutter free throughout the expanded design envelope. All the physical parameters that have been determined to be significant for flutter response should be appropriately scaled. These will include elastic and inertia properties, geometric properties, and dynamic pressure. If transonic effects are important, the Mach number should be maintained.
7.2.4.3 Validation of analysis methods is another appropriate use of wind tunnel flutter testing. When the validity of a method is uncertain, correlation of wind tunnel flutter testing results with a corresponding analysis may increase confidence in the use of the analytical tool for certification analysis. A methods validation test should simulate conditions, scaling, and geometry appropriate for the intended use of the analytical method.
7.2.4.4 Trend studies are an important use of wind tunnel flutter testing.
Parametric studies can be used to establish trends for control system balance and stiffness, fuel and payload variations, structural compliances, and configuration variations. The set of physical parameters requiring similitude may not be as extensive to study parametric trends as is necessary for clearance testing. For example, an exact match of the Mach number may not be needed to track the effects of payload variations on a transonic airplane.
7.2.5 Flight Flutter Tests.
7.2.5.1 Full-scale flight flutter testing of an airplane configuration to VDF/MDF is a necessary part of the flutter substantiation. An exception may be made when aerodynamic, mass, or stiffness changes to a certified airplane are minor, and analysis or ground tests show a negligible effect on flutter or vibration characteristics. If a failure, malfunction, or adverse condition is simulated during a flight test, the maximum speed investigated need not exceed VFC/MFC if it is shown, by correlation of the flight test data with other test data or analyses, that the requirements of § 25.629(b)(2) are met.
7.2.5.2 Airplane configurations and control system configurations should be selected for flight test based on analyses and, when available, model test results. Sufficient test conditions should be performed to demonstrate aeroelastic stability throughout the entire flight envelope for the selected configurations.
7.2.5.3 To achieve adequate test results, flight flutter testing requires excitation sufficient to excite the modes shown by analysis to be the most likely to couple for flutter. Excitation methods may include control surface motions or internal moving mass or external aerodynamic exciters or flight turbulence. Use the appropriate method of excitation for the modal 19
AC 25.629-1C
response frequency being investigated. The effect of the excitation system itself on the airplane flutter characteristics should be determined prior to flight testing.
7.2.5.4 Measurement of the response at selected locations on the structure should be made to determine the response amplitude, damping, and frequency in the critical modes at each test airspeed. It is desirable to monitor the response amplitude, frequency, and damping change as VDF/MDF is approached. In demonstrating that there is no large and rapid damping reduction as VDF/MDF is approached, an endeavor should be made to identify a clear trend of damping versus speed. If this is not possible, then sufficient test points should be undertaken to achieve a satisfactory level of confidence that there is no evidence of an adverse trend.
7.2.5.5 An evaluation of phenomena not presently amenable to analyses, such as shock effects, buffet response levels, vibration levels, and control surface buzz, should also be made during flight testing.
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