There are other requirements. Are you a FAA approved mechanic or are you doing this on your our airplane? Do you have repairmens certificate?
U.S. Department
of Transportation
Federal Aviation
Administration
Advisory
Circular
Subject: AEROELASTIC STABILITY Date: 7/23/98 AC No: 25.629-1A
SUBSTANTIATION OF TRANSPORT Initiated by: ANM- 114 Change:
CATEGORY AIRPLANES
1. PURPOSE. This advisory circular (AC) provides guidance material for
acceptable means, but not the only means, of demonstrating compliance with the
provisions of part 25 of the Federal Aviation Regulations (FAR) dealing with the design
requirements for transport category airplanes to preclude the aeroelastic instabilities of
flutter, divergence and control reversal. The precise detailing of analytical procedures
and testing techniques is beyond the scope of this AC. Some general considerations are
set forth herein, with supportive discussion, to be considered in demonstrating
compliance with § 25.629 and related regulations.
2. CANCELLATION. Advisory Circular 25.629-1, Flutter Substantiation of
Transport Category Airplanes, dated January 4, 1985, is cancelled.
3. RELATED FAR SECTIONS.
§ 25.251 - Vibration and buffeting.
§ 25.305 - Strength and deformation.
§ 25.335 - Design airspeeds.
§ 25.343 - Design fuel and oil loads.
§ 25.571 - Damage-tolerance and fatigue evaluation of structure.
§ 25.629 - Aeroelastic stability requirements.
§ 25.631 - Bird strike damage.
§ 25.671 - General (Control Systems).
§ 25.672 - Stability augmentation and automatic and power-operated systems.
§ 25.1309 - Equipment, systems and installations.
§ 25.1329 - Automatic pilot system.
§ 25.1419 - Ice protection.
4. BACKGROUND.
a. Flutter and other aeroelastic instability phenomena have had a significant influence on airplane
development and the airworthiness criteria governing the design of civil airplanes. The initial requirement for
consideration of flutter was minimal in the
FAA Form 1320-15 (4-82) Supersedes WA Form 1320-2
AC 25.629-1A 7/23/98
1931 "Airworthiness Requirements of Air Commercial Regulations for Aircraft," Bulletin No. 7-A. The airplane
flutter requirement specified that "no surface shall show any signs of flutter or appreciable vibration in any
attitude or condition of flight." In 1934, Bulletin No. 7-A was revised in view of service experience and
contained advice and good practice techniques for the early airplane designer regarding flutter prevention
measures. All airplane designs were required to have interconnected elevators, statically-balanced ailerons,
irreversible or balanced tabs, and, in some cases, a ground vibration test was required to be conducted.
b. Regulations dealing specifically with flutter, deformation, and vibration on transport category
airplanes were first- introduced when part 04 of the Civil Air Regulations (CAR) became effective in the mid- I
940's. The criteria related the solution of the flutter problem to frequency ratios based on model tests conducted
by the Army Air Corps. Also, based on the Army Air Corps developments, part 04 imposed a design factor of
1.2 on equivalent airspeed to provide a stiffness margin for the airframe. In addition to this empirical approach,
and recognizing the advancing state-of-the-art, part 04 referenced publications containing developing flutter
theory.
C. The flutter requirement of part 04 evolved into CAR 4b.308 where developing fail-safe
philosophy continued to change the scope of flutter substantiation. Among these developments was a revision to
CAR 4b.320 in 1956 to require fail-safe tabs and a revision to CAR 4b.308 in 1959 to require fail-safe flutter
damper installations. The flutter requirement was extensively revised in 1964 to require compliance with the
single failure criteria for the entire airplane as well as adding special provisions for turboprop airplanes.
d. Service experience indicated that single failure criteria related to flutter stability were not
sufficiently objective and comprehensive to cover modem, complex, transport airplanes equipped with highly
redundant systems. Therefore, part 25 of the FAR, which was recodified from part 04b of the CAR, was
amended to require that, unless combinations of failures are shown to be extremely improbable, they must be
considered in design for freedom from flutter and divergence.
e. The development of speed and attitude limiting systems has created the need for a minimum
speed margin for fail-safe aeroelastic stability substantiation. Part 25 as amended at Amendment 77
incorporated this minimum fail-safe speed boundary, revised the safety margins for aeroelastic stability, and
expanded the list of failures, malfunctions and adverse conditions that needed to be addressed.
f. Additional regulations governing the interaction of systems with structures have been written
for airplanes with advanced electronic flight control systems. These regulations prescribe variations in the
fail-safe speed margins depending on the probability of system failure.
5. . DISCUSSION OF REQUIREMENTS. The general requirement for demonstrating freedom from aeroelastic
instability is contained in § 25.629, which also sets forth specific requirements for the investigation of these aeroelastic
phenomena for various airplane configurations and flight conditions. Additionally, there are other conditions defined by
the sections of the FAR listed in paragraph 3 above to be investigated for aeroelastic stability to assure safe flight.
Many of the conditions contained in this AC pertain only to certain specific amendments of the FAR. Type design
changes to airplanes certified to an earlier part 25 amendment must meet the certification basis established for the
modified airplane.
a. Aeroelastic Stability Envelope.
(1) For nominal conditions without failures, malfunctions, or adverse conditions, freedom from
aeroelastic instability is required to be shown for all combinations of airspeed and altitude encompassed by the design
dive speed (VD) and design dive Mach number (MD) versus altitude envelope enlarged at all points by an increase of
15 percent in equivalent airspeed at both constant Mach number and constant altitude. Figure I A represents a typical
design envelope expanded to the required aeroelastic stability envelope. Note that some required Mach number and
airspeed combinations correspond to altitudes below standard sea level.
(2) The aeroelastic stability envelope may be limited to a maximum Mach number of 1.0 when
MD is less than 1.0 and when there is no large and rapid reduction in damping as MD is approached.
(3) Some configurations and conditions that are required to be investigated by § 25.629 and other
part 25 regulations consist of failures, malfunctions or adverse conditions. Aeroelastic stability investigations of these
fail-safe conditions need to be carried out for all approved altitudes to the greater airspeed defined by:
(a) The V
D/
M
D envelope determined by § 25.335(b); or,
(b) An altitude-airspeed envelope defined by a 15 percent increase in equivalent airspeed
above V
C at constant altitude, from sea level up to the altitude of the intersection of 1. 15 V
C with the extension of the
constant cruise Mach number line, MC, then a linear variation in equivalent airspeed to M
C + .05 at the altitude of the
lowest VC/MC intersection; then at higher altitudes, up to the maximum flight altitude, the boundary defined by a .05
Mach increase in MC at constant altitude.
Figure IB shows the minimum aeroelastic stability envelope for fail-safe conditions, which is a composite of the
highest speed at each altitude from either the VD envelope or the constructed altitude-airspeed envelope based on the
defined V
C and M
C.
Fail-safe design speeds, other than the ones defined above, may be used for certain system failure conditions
when specifically authorized by other rules or special conditions prescribed in the certification basis of the airplane.
b. Configurations and Conditions. The following paragraphs provide a summary of the configurations
and conditions to be investigated in demonstrating compliance with part 25. Specific design configurations may
warrant additional considerations not discussed in this AC.
(1) Nominal Configurations and Conditions. Nominal configurations and conditions of the
airplane are those that are likely to exist during normal operation. Freedom from aeroelastic instability should be
shown throughout the expanded clearance envelope described in paragraph 5a(l) above for:
(a) The range of fuel and payload combinations, including zero fuel, for which
certification is requested.
(b) Configurations with any likely ice mass accumulations on unprotected surfaces for
airplanes approved for operation in icing conditions.
(c) All normal combinations of autopilot, yaw damper, or other automatic flight control
systems.
(d) All possible engine settings and combinations of settings from idle power to
maximum available thrust including the conditions of one engine stopped and windmilling, in order to address the
influence of gyroscopic loads and thrust on aeroelastic stability.
(2) Failures, Malfunctions. and Adverse Conditions. The following conditions should be investigated
for aeroelastic instability within the fail-safe envelope defined in paragraph 5a(3) above.
(a) Any critical fuel loading conditions, not shown to be extremely improbable, which
may result from mismanagement of fuel.
(b) Any single failure in any flutter control system.
(c) For airplanes not approved for operation in icing conditions, any likely ice
accumulation expected as a result of an inadvertent encounter. For airplanes approved for operation in icing
conditions, the maximum likely ice accumulation expected as the result of any single failure in the de-icing system, or
any combination of failures not shown to be extremely improbable.
(d) Failure of any single element of the structure supporting any engine, independently
mounted propeller shaft, large auxiliary power unit, or large externally mounted aerodynamic body (such as an
external fuel tank).
(e) For airplanes with engines that have propellers or large rotating devices capable of
significant dynamic forces, any single failure of the engine structure that would reduce the rigidity of the rotational
axis.
(f) The absence of aerodynamic or gyroscopic forces resulting from the most adverse
combination of feathered propellers or other rotating devices capable of significant dynamic forces. In addition, the
effect of a single feathered propeller or rotating device must be coupled with the failures of paragraphs 5b(2)(d) and
5b(2)(e) above.
(g) Any single propeller or rotating device capable of significant dynamic forces rotating
at the highest likely overspeed.
(h) Any damage or failure condition, required or selected for investigation by § 25.571. The
single structural failures described in paragraphs 5b(2)(d) and 5b(2)(e) above need not be considered in showing
compliance with this paragraph if;
1 The structural element could not fail due to discrete source damage resulting
from the conditions described in § 25.571(e); and
2 A damage tolerance investigation in accordance with § 25.571(b) shows that
the maximum extent of damage assumed for the purpose of residual strength evaluation does not involve complete
failure of the structural element.
(i) Any damage, failure or malfunction, considered under §§ 25.631, 25.671, 25.672,
and 25.1309. This includes the condition of two or more engines stopped or windmilling for the design range of fuel
and payload combinations, including zero fuel.
(j) Any other combination of failures, malfunctions, or adverse conditions not shown to
be extremely improbable.
C. Detail Design Requirements.
(1) Main surfaces, such as wings and stabilizers, should be designed to meet the aeroelastic
stability criteria for nominal conditions and should be investigated for meeting fail-safe criteria by considering stiffness
changes due to discrete damage or by reasonable parametric variations of design values.
(2) Control surfaces, including tabs, should be investigated for nominal conditions and for failure
modes that include single structural failures (such as actuator disconnects, hinge failures, or, in the case of
aerodynamic balance panels, failed seals), single and dual hydraulic system failures and any other combination of
failures not shown to be extremely improbable. Where other structural components contribute to the aeroelastic
stability of the system, failures of those components should be considered for possible adverse effects.
(3) Where aeroelastic stability relies on control system stiffness and/or damping, additional
conditions should be considered. The actuation system should continuously provide, at least, the minimum stiffness or
damping required for showing aeroelastic stability without regard to probability of occurrence for:
(a) More than one engine stopped or windmilling,
(b) Any discrete single failure resulting in a change of the
structural modes of vibration (for example; a disconnect or failure of a mechanical element, or a structural failure of a
hydraulic element, such as a hydraulic line, an actuator, a spool housing or a valve),
(c) Any damage or failure conditions considered under §§ 25.571, 25.631 and 25.671.
The actuation system minimum requirements should also be continuously met after any combination of failures
not shown to be extremely improbable (occurrence less than 10-9 per flight hour). However, certain combinations of
failures, such as dual electric or dual hydraulic system failures, or any single failure in combination with any probable
electric or hydraulic system failure (§ 25.671), are not normally considered extremely improbable regardless of
probability calculations. The reliability assessment should be part of the substantiation documentation. In practice,
meeting the above conditions may involve design concepts such as the use of check valves and accumulators,
computerized pre-flight system checks and shortened inspection intervals to protect against undetected failures.
(4) Consideration of free play may be incorporated as a variation in stiffness to assure adequate
limits are established for wear of components such as control surface actuators, hinge bearings, and engine mounts in
order to maintain aeroelastic stability margins.
(5) If balance weights are used on control surfaces, their effectiveness and strength, including
that of their support structure, should be substantiated.
(6) The automatic flight control system should not interact with the airframe to produce an
aeroelastic instability. When analyses indicate possible adverse coupling, tests should be performed to determine the
dynamic characteristics of actuation systems such as servo-boost, fully powered servo-control systems, closed-loop
airplane flight control systems, stability augmentation systems, and other related powered-control systems.
6. COMPLIANCE. Demonstration of compliance with aeroelastic stability requirements for an airplane
configuration may be shown by analyses, tests, or some combination thereof. In most instances, analyses are required
to determine aeroelastic stability margins for normal operations, as well as for possible failure conditions. Wind tunnel
flutter model tests, where applicable, may be used to supplement flutter analyses. Ground testing may be used to collect
stiffness or modal data for the airplane or components. Flight testing may be used to demonstrate compliance of the
airplane design throughout the design speed envelope.
a. Analytical Investigations. Analyses should normally be used to investigate the aeroelastic stability of
the airplane throughout its design flight envelope and as expanded by the required speed margins. Analyses are used to
evaluate aeroelastic stability sensitive parameters such as aerodynamic coefficients, stiffness and mass distributions,
control surface balance requirements, fuel management schedules, engine/store locations, and control system
characteristics. The sensitivity of most critical parameters may be determined analytically by varying the parameters
from nominal. These investigations are an effective way to account for the operating conditions and possible failure
modes which may have an effect on aeroelastic stability margins, and to account for uncertainties in the values of
parameters and expected variations due to in-service wear or failure conditions.
(1) Analytical Modeling. The following sections discuss acceptable, but not the only, methods and
forms of modeling airplane configurations and/or components for purposes of aeroelastic stability analysis. The types of
investigations generally encountered in the course of airplane aeroelastic stability substantiation are also discussed. The
basic elements to be modelled in aeroelastic stability analyses are the elastic, inertial, and aerodynamic characteristics
of the system. The degree of complexity required in the modeling, and the degree to which other characteristics need
to be included in the modeling, depend upon the system complexity.
(a) Structural Modeling. Most forms of structural modeling can be classified into two
main categories: (1) modeling using a lumped mass beam, and (2) finite element modeling. Regardless of the approach
taken for structural modeling, a minimum acceptable level of sophistication, consistent with configuration complexity, is
necessary to satisfactorily represent the critical modes of deformation of the primary structure and control surfaces.
The model should reflect the support structure for the attachment of control surface actuators, flutter dampers, and any
other elements for which stiffness is important in prevention of aeroelastic instability. Wing-pylon mounted engines are
often significant to aeroelastic stability and warrant particular attention in the modeling of the pylon, and pylon-engine
and pylon-wing interfaces. The model should include the effects of cut-outs, doors, and other structural features which
may tend to affect the resulting structural effectiveness. Reduced stiffness should be considered in the modeling of
airplane structural components which may exhibit some change in stiffness under limit design flight conditions.
Structural models include mass distributions as well as representations of stiffness and possibly damping
characteristics. Results from the models should be compared to test data, such as that obtained from ground vibration
tests, in order to determine the accuracy of the model and its applicability to the aeroelastic stability investigation.
(b) Aerodynamic Modeling.
1 Aerodynamic modeling for aeroelastic stability requires the use of unsteady,
two-dimensional strip or three-dimensional panel theory methods for incompressible or compressible flow. The choice
of the appropriate technique depends on the complexity of the dynamic structural motion of the surfaces under
investigation and the flight speed envelope of the airplane. Aerodynamic modeling should be supported by tests or
previous experience with applications to similar configurations.
2 Main and control surface aerodynamic data are commonly adjusted by weighting
factors in the aeroelastic stability solutions. The weighting factors for steady flow (k=O) are usually obtained by
comparing wind tunnel test results with theoretical data. Special attention should be given to control surface
aerodynamics because viscous and other effects may require more extensive adjustments to theoretical coefficients.
Main surface aerodynamic loading due to control surface deflection should be considered.
(2) Types of Analyses.
(a) Oscillatory (flutter) and non-oscillatory (divergence and control reversal) aeroelastic
instabilities should be analyzed to show compliance with § 25.629.
(b) 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.
(c) 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.
(d) 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.
(3) Damping Requirements.
(a) There is no intent in this AC to define a flight test level of acceptable minimum
damping.
(b) Flutter analyses results are usually presented graphically in the form of frequency
versus velocity (V-f, Figure 2) and damping versus velocity (V-g, Figures 3 and 4) curves for each root of the flutter
solution.
(c) Figure 3 details one common method for showing compliance with the requirement
for a proper margin of damping. It is based on the assumption that the structural damping available is 0.03 (1.5%
critical viscous damping) and is the same for all modes as depicted by the V-g curves shown in Figure 3. No significant
mode, such as curves (2) or (4), should cross the g=O 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 3, 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 L 15 VD-
(d) Figure 4 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=O line at zero airspeed and, in
some cases, flutter solutions different from those obtained with no structural damping. The similarity in the curves of
Figures 3 and 4 are only for simplifying this example. The minimum acceptable damping line applied to the analytical
results as shown in Figure 4 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=O line below 1. 15 VD.
(e) 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 5a(3) above.
(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, and 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 which
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 free play, and control system
components which 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.
(a) 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 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 analysed 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.
(b) Mass Balance.
1 The magnitude and spanwise location of control surface balance weights may be
evaluated 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 must be maintained between the frequency of the control surface
first torsion mode and the flutter mode.
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.
Free play between the balance weight, the support arm, and the control surface must not be allowed. Control surface
mass properties (weight and static unbalance) should be confirmed by measurement before ground vibration testing.
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
(c) 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 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 free play 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. Refer to paragraph 5c(3) above 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.
(d) 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 which 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.
(e) Ice Accumulation. Aeroelastic stability analysis should use the mass distributions
derived from the maximum likely ice accumulations. The ice accumulation determination can take into account the
ability to detect the ice and the time required to leave the icing condition. The analyses need not consider the
aerodynamic effects of ice shapes.
(f) Whirl Flutter.
1 The 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.
2 Failure conditions are usually significant for whirl instabilities. Engine mount,
engine gear box support, or shaft failures which result in a node line shift for propeller hub pitching or yawing motion
are especially significant.
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.
(g) Automatic Control Systems. 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 which significantly affect the system gain and/or phase and are not shown to be extremely
improbable should be analysed.
b. 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.
(1) Structural Component Tests. Stiffness tests or ground vibration
tests 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.
(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 tests are performed off the airplane
the test fixtures should reflect local attachment flexibility.
(a) 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.
(b) 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, however, 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 free play should be measured.
(3) Ground Vibration Tests.
(a) Ground vibration tests (GVT) 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 modified to match test
data, may be acceptable provided 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.
(b) '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 ground vibration tests and is necessary in
the absence of vibration analysis.
(c) 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 insure 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.
(d) 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 accurately
describe the response or mode shapes of all significant structural modes. In addition, the structural damping of each
mode should be determined.
(4) Flutter Model Tests.
(a) 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.
(b) 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 which 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.
(c) 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.
(d) 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 required for clearance testing. For example, an exact match of the Mach number may
not be required to track the effects of payload variations on a transonic airplane.
(5) Flight Flutter Tests.
(a) 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.
(b) 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.
(c) 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. The method of excitation must be appropriate for
the modal response frequency being investigated. The effect of the excitation system itself on the airplane flutter
characteristics should be determined prior to flight testing.
(d) Measurement of the response at selected locations on the structure should be made in
order 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.
(e) 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.