0 

U.S. Department 

of Transportation 

Federal Aviation 

Administration 

Advisory


Circular 

Subject: Aeroelastic Stability Substantiation of 

Transport Category Airplanes 

Date: 10/27/14 ACNo:25.629-1B 

Initiated By: ANM-115 

This advisory circular (AC) provides guidance material for acceptable means, but not the only 

means, of demonstrating compliance with the provisions of Title 14, Code of Federal 

Regulations part 25 associated with the design requirements for transport category airplanes to 

preclude the aeroelastic instabilities of flutter, divergence, and control reversal. The precise 

details for analytical procedures and testing techniques are beyond the scope of this AC. 

However, some general information is included, with supporting discussions, for consideration 

when demonstrating compliance with§ 25.629 and related regulations. Revision B incorporates 

changes only to the portions of this AC affected by the Airplane and Engine Certification 

Requirements in Supercooled Large Drop, Mixed Phase, and Ice Crystal Icing Conditions final 

rule. 

If you have suggestions for improving this AC, you may use the feedback form at the end ofthis 

AC. 

~~Michael Kaszycki 

Acting Manager, Transport Airplane Directorate 

Aircraft Certification Service 
10/27/14 AC 25.629-1B

Contents

Paragraph Page

ii

1 Purpose.......................................................................................................................................1

2 Applicability. .............................................................................................................................1

3 Cancellation. ..............................................................................................................................1

4 Related 14 CFR Regulations......................................................................................................1

5 Background................................................................................................................................2

6 Discussion of Requirements. .....................................................................................................3

6.1 Aeroelastic Stability Envelope..........................................................................................3

6.2 Configurations and Conditions. ........................................................................................5

6.2.1 Nominal Configurations and Conditions. .............................................................5

6.2.2 Failures, Malfunctions, and Adverse Conditions..................................................6

6.3 Detail Design Requirements. ............................................................................................7

7 Compliance. ...............................................................................................................................8

7.1 Analytical Investigations. .................................................................................................8

7.1.1 Analytical Modeling. ............................................................................................9

7.1.2 Types of Analyses...............................................................................................10

7.1.3 Damping Requirements. .....................................................................................10

7.1.4 Analysis Considerations......................................................................................12

7.2 Testing.............................................................................................................................15

7.2.1 Structural Component Tests................................................................................16

7.2.2 Control System Component Tests. .....................................................................16

7.2.3 Ground Vibration Tests.......................................................................................16

7.2.4 Flutter Model Tests.............................................................................................17

7.2.5 Flight Flutter Tests..............................................................................................18
10/27/14 AC 25.629-1C

Contents (continued)

iii

Figures

Number Page

Figure 1. Minimum Required Aeroelastic Stability Margin........................................................... 4

Figure 2. Minimum Fail-Safe Clearance Envelope ........................................................................ 5

Figure 3. Frequency versus Velocity ............................................................................................ 11

Figure 4. Damping versus Velocity—Method 1........................................................................... 11

Figure 5. Damping versus Velocity—Method 2........................................................................... 12
10/27/14 AC 25.629-1B

1

1 PURPOSE.

This AC provides guidance material for acceptable means, but not the only means, of 

demonstrating compliance with the provisions of Title 14, Code of Federal Regulations 

(14 CFR) part 25 associated with the design requirements for transport category 

airplanes to preclude the aeroelastic instabilities of flutter, divergence, and control 

reversal. The precise details for analytical procedures and testing techniques are beyond 

the scope of this AC. However, some general information is included, with supporting 

discussions, for consideration when demonstrating compliance with § 25.629 and 

related regulations. Revision B incorporates changes to only the portions of this AC 

affected by the Airplane and Engine Certification Requirements in Supercooled Large 

Drop, Mixed Phase, and Ice Crystal Icing Conditions final rule (79 FR 65508, 

November 4, 2014).

2 APPLICABILITY.

2.1 The guidance provided in this document is directed to airplane manufacturers, 

modifiers, foreign regulatory authorities, and Federal Aviation Administration (FAA) 

transport airplane type certification engineers and their designees.

2.2 The material in this AC is neither mandatory nor regulatory in nature and does not 

constitute a regulation. It describes acceptable means, but not the only means, for 

demonstrating compliance with the applicable regulations. The FAA will consider other 

means of demonstrating compliance that an applicant may elect to present. While these 

guidelines are not mandatory, they are derived from extensive FAA and industry 

experience in determining compliance with the relevant regulations. On the other hand, 

if we become aware of circumstances that convince us that following this AC would not 

result in compliance with the applicable regulations, we will not be bound by the terms 

of this AC, and we may require additional substantiation or design changes as a basis 

for finding compliance.

2.3 The material in this AC does not change, create any additional, authorize changes in, or 

permit deviations from, regulatory requirements.

3 CANCELLATION.

This AC cancels AC 25.629-1A, Aeroelastic Stability Substantiation of Transport 

Category Airplanes, dated July 23, 1998.

4 RELATED 14 CFR REGULATIONS.

• Section 25.251, Vibration and buffeting.

• Section 25.305, Strength and deformation.

• Section 25.335, Design airspeeds.

• Section 25.343, Design fuel and oil loads.
10/27/14 AC 25.629-1B

2

• Section 25.571, Damage-tolerance and fatigue evaluation of structure.

• Section 25.629, Aeroelastic stability requirements.

• Section 25.631, Bird strike damage.

• Section 25.671, General (Control Systems).

• Section 25.672, Stability augmentation and automatic and power-operated systems.

• Section 25.1309, Equipment, systems and installations.

• Section 25.1329, Flight guidance system.

• Section 25.1419, Ice protection.

• Section 25.1420, Airplane and Engine Certification Requirements in Supercooled 

Large Drop, Mixed Phase, and Ice Crystal Icing Conditions.

5 BACKGROUND.

5.1 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 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.

5.2 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-1940s. 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.

5.3 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.
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3

5.4 Service experience indicated that single failure criteria related to flutter stability were 

not sufficiently objective and comprehensive to cover modern, complex, transport 

airplanes equipped with highly redundant systems. Therefore, part 25, 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.

5.5 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 by Amendment 25-77 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.

5.6 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.

6 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 part 25 

listed in paragraph 4, 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 part 25. Type design changes to airplanes certified to an earlier part 25 

amendment must meet the certification basis established for the modified airplane.

6.1 Aeroelastic Stability Envelope.

6.1.1 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 1

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.
10/27/14 AC 25.629-1B

4

Figure 1. Minimum Required Aeroelastic Stability Margin

VE

Equivalent

Airspeed

Sea

Level

15% VE

Constant

Mach No.

15% VE Constant

Altitude

Typical VD/MD Envelope

Maximum Design Altitude

C” C’

B’

A’

C B

A

MD

Mach Number

VD

6.1.2 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.

6.1.3 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.

6.1.3.1 Aeroelastic stability investigations of these fail-safe conditions need to be 

carried out for all approved altitudes to the greater airspeed defined by—

6.1.3.1.1 The VD/MD envelope determined by § 25.335(b); or

6.1.3.1.2 An altitude-airspeed envelope defined by a 15 percent increase in 

equivalent airspeed above VC at constant altitude, from sea level up to the 

altitude of the intersection of 1.15 VC with the extension of the constant 

cruise Mach number line, MC, then a linear variation in equivalent 

airspeed to MC + 0.05 at the altitude of the lowest VC/MC intersection; 

then at higher altitudes, up to the maximum flight altitude, the boundary 

defined by a 0.05 Mach increase in MC at constant altitude.

6.1.3.2 Figure 2 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 VC and MC.
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6.1.3.3 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.

Figure 2. Minimum Fail-Safe Clearance Envelope

Altitude

Airspeed

1.15 VC

MD

VD

MC + .05

6.2 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.

6.2.1 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 6.1.1 above for the 

following:

6.2.1.1 The range of fuel and payload combinations, including zero fuel, for 

which certification is requested.

6.2.1.2 Configurations with ice mass accumulations on unprotected surfaces for 

airplanes approved for operation in icing conditions. See paragraph 7.1.4.5

of this AC.

6.2.1.3 All normal combinations of autopilot, yaw damper, or other automatic 

flight control systems.
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6

6.2.1.4 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.

6.2.2 Failures, Malfunctions, and Adverse Conditions.

The following conditions should be investigated for aeroelastic instability within the 

fail-safe envelope defined in paragraph 6.1.3 above.

6.2.2.1 Any critical fuel loading conditions, not shown to be extremely 

improbable, which may result from mismanagement of fuel.

6.2.2.2 Any single failure in any flutter control system.

6.2.2.3 For airplanes not approved for operation in icing conditions, ice 

accumulation expected as a result of an inadvertent encounter. For 

airplanes approved for operation in icing conditions, the 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. See 

paragraph 7.1.4.5 of this AC.

6.2.2.4 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).

6.2.2.5 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.

6.2.2.6 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 in 

paragraphs 6.2.2.4 and 6.2.2.5 above.

6.2.2.7 Any single propeller or rotating device capable of significant dynamic 

forces rotating at the highest likely overspeed.

6.2.2.8 Any damage or failure condition, required or selected for investigation by 

§ 25.571. The single structural failures described in paragraphs 6.2.2.4 and 

6.2.2.5 above need not be considered in showing compliance with this 

paragraph if—

6.2.2.8.1 The structural element could not fail due to discrete source damage 

resulting from the conditions described in § 25.571(e); and

6.2.2.8.2 A damage tolerance investigation in accordance with § 25.571(b) shows 

that the maximum extent of damage assumed for the purpose of residual 
10/27/14 AC 25.629-1B

7

strength evaluation does not involve complete failure of the structural 

element.

6.2.2.9 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.

6.2.2.10 Any other combination of failures, malfunctions, or adverse conditions not 

shown to be extremely improbable.

6.3 Detail Design Requirements.

6.3.1 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.

6.3.2 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.

6.3.3 Where aeroelastic stability relies on control system stiffness and/or damping, additional 

conditions should be considered.

6.3.3.1 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—

6.3.3.1.1 More than one engine stopped or windmilling;

6.3.3.1.2 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); and

6.3.3.1.3 Any damage or failure conditions considered under §§ 25.571, 25.631, and 

25.671.

6.3.3.2 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 
10/27/14 AC 25.629-1B

8

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.

6.3.4 Consideration of freeplay 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.

6.3.5 If balance weights are used on control surfaces, their effectiveness and strength, 

including that of their support structure, should be substantiated.

6.3.6 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.

7 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.

7.1 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 that

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.
10/27/14 AC 25.629-1B

9

7.1.1 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 

modeled 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.

7.1.1.1 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 that may tend to affect the resulting structural 

effectiveness. Reduced stiffness should be considered in the modeling of 

airplane structural components that 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.

7.1.1.2 Aerodynamic Modeling.

7.1.1.2.1 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.

7.1.1.2.2 Main and control surface aerodynamic data are commonly adjusted by 

weighting factors in the aeroelastic stability solutions. The weighting 

factors for steady flow (k=0) are usually obtained by comparing wind 

tunnel test results with theoretical data. Special attention should be given 
10/27/14 AC 25.629-1B

10

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.

7.1.2 Types of Analyses.

7.1.2.1 Oscillatory (flutter) and non-oscillatory (divergence and control reversal) 

aeroelastic instabilities should be analyzed to show compliance with 

§ 25.629.

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.
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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 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 4. No 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
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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.

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 above.

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, 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 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 
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13

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 

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 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 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 ground vibration testing.
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14

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 

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

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.

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 part 25, Appendices C and O. This includes 
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15

any accretions that could develop on control surfaces. The analyses need 

not consider the aerodynamic effects of ice shapes. For airplanes approved 

for operation in icing conditions, all of 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 of the Appendix C and O 

icing conditions 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.

7.1.4.6 Whirl Flutter.

7.1.4.6.1 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.

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.

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.

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 
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16

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 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.

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 tests are performed 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, 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 freeplay should be measured.

7.2.3 Ground Vibration Tests.

7.2.3.1 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.

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 
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17

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.

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 accurately describe 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.

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.
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18

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 

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.

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 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 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 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 
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19

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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