AC 25.629-1C

U.S. Department of Transportation Federal Aviation Administration Advisory Circular Subject: Aeroelastic Stability Substantiation of Transport Category Airplanes 08/30/2024Date: AC No: 25.629-1C Initiated By: AIR-600 DANIEL J. Digitally signed by DANIEL J. ELGAS ELGAS Date: 2024.08.30 21:10:42 -04'00' Daniel Elgas Director, Policy and Standards Division Aircraft Certification Service

This advisory circular (AC) provides general guidance for acceptable means, but not the only means, of demonstrating compliance with those provisions of Title 14, Code of Federal Regulations (14 CFR) 25.629, Aeroelastic stability requirements, and other part 25 regulations related to aeroelastic instabilities of flutter, divergence, and control reversal. Revision C of this AC contains guidance based on rule changes to § 25.629.

If you have suggestions for improving this AC, you may use the Advisory Circular Feedback form at the end of this AC.

AC 25.629-1C

CONTENTS Paragraph Page

1 Purpose1

2 Applicability1

3 Cancellation1

4 Related 14 CFR Regulations1

5 Background2

6 Discussion of Requirements3

6.1 Aeroelastic Stability Envelope3

6.2 Configurations and Conditions5

7 Compliance9

7.1 Analytical Investigations9

7.2 Testing17

FIGURES Number Page

Figure 1. Minimum Required Aeroelastic Stability Margin4

Figure 2. Minimum Fail-Safe Clearance Envelope5

Figure 3. Frequency versus Velocity11

Figure 4. Damping versus Velocity—Method 112

Figure 5. Damping versus Velocity—Method 213

ii

AC 25.629-1C

1 PURPOSE.

This AC provides guidance for acceptable means, but not the only means, of demonstrating compliance with 14 CFR 25.629, and other provisions of part 25 intended to preclude the aeroelastic instabilities of flutter, divergence, and control reversal. The details of all possible analytical procedures and testing techniques are beyond the scope of this AC. However, this AC includes general information for applicants to consider when demonstrating compliance with § 25.629 and related regulations. Revision C of this AC contains guidance based on rule changes to § 25.629.

The contents of this document do not have the force and effect of law and are not meant to bind the public in any way. This document is intended only to provide clarity to the public regarding existing requirements under the law or agency policies.

2 APPLICABILITY.

2.1 The guidance provided in this AC is for airplane manufacturers, modifiers, foreign regulatory authorities, and Federal Aviation Administration (FAA) transport airplane type certification engineers and their designees.

2.2 Using this guidance as a means of compliance to § 25.629 is voluntary only and not using it will not affect rights and obligations under existing statutes and regulations. The FAA will consider other methods of demonstrating compliance that an applicant may elect to present. Terms such as “should,” “may,” and “must” are used only in the sense of ensuring applicability of this particular method of compliance when the acceptable method of compliance in this document is used. If the FAA becomes aware of circumstances in which following this AC would not result in compliance with the applicable regulations, the agency may require additional substantiation as the basis for finding compliance.

3 CANCELLATION.

This AC cancels AC 25.629-1B, Aeroelastic Stability Substantiation of Transport Category Airplanes, dated October 27, 2014.

4 RELATED 14 CFR REGULATIONS.

The following 14 CFR part 25 regulations are related to this AC. You can download the full text of these regulations from the Federal Register website at Electronic Code of Federal Regulations, jointly administered by the Office of the Federal Register (OFR) of the National Archives and Records Administration (NARA) and the U.S.

Government Publishing Office (GPO). You can order a paper copy from the U.S.

Superintendent of Documents, U.S. Government Publishing Office, Washington, D.C.

20401; at Government Publishing Office, by calling telephone number (202) 512-1800;

or by sending a fax to (202) 512-2250.

• Section 25.251, Vibration and buffeting.

• Section 25.305, Strength and deformation.

AC 25.629-1C

• Section 25.335, Design airspeeds.

• Section 25.343, Design fuel and oil loads.

• 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, Control Systems—General.

• 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, Supercooled large drop 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 brief 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, the U.S. Civil Aeronautics Board (CAB) revised Bulletin No. 7-A in view of service experience, providing advice and good practice techniques for the early airplane designer regarding flutter prevention measures. This revision also required all airplane designs to have interconnected elevators, statically-balanced ailerons, irreversible or balanced tabs, and, in some cases, a ground vibration test (GVT).

5.2 The CAB introduced regulations on flutter, deformation, and vibration on transport category airplanes in the mid-1940s, through various provisions of part 04 of the Civil Air Regulations (CAR). The design criteria related the solution of the flutter problem to frequency ratios that were based on model tests conducted by the Army Air Corps.

Also, based on the Army Air Corps’ research, part 04 imposed a safety factor of 1.2 on the required equivalent airspeed, to provide a stiffness margin for the airframe. Since aircraft technology was rapidly changing, part 04 also referenced publications containing flutter theory.

5.3 The flutter requirement of part 04 evolved into CAR 4b.308, while developing fail-safe engineering philosophy continued to change the scope of flutter substantiation. Fail-safe design is one in which a single failure would not adversely affect safety of flight.

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 FAA extensively revised the flutter requirement in 1964 to require 2

AC 25.629-1C that the entire airplane comply with the single-failure criteria, and to add special provisions for turboprop airplanes.

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, the FAA amended part 25, the successor to part 04b of the CAR, to require that, unless an applicant could show that combinations of failures were extremely improbable, the applicant must consider such combinations when designing the airplane to be free from flutter and divergence.

5.5 The development of speed and attitude limiting systems 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 The effect of flight control systems, autopilots, stability augmentation systems, load alleviation systems, and other systems that can affect aeroelastic and aeroservoelastic stability under nominal conditions is addressed in terms of gain and phase margins that characterize feedback control systems. In addition, clarifications were added about assessing the effect of high load factors on aeroelastic stability.

6 DISCUSSION OF REQUIREMENTS.

The general requirement for demonstrating freedom from aeroelastic instability is in § 25.629, which also requires investigation of these aeroelastic phenomena for various airplane configurations and flight conditions. Additionally, to assure safe flight, the applicant must investigate other conditions for aeroelastic stability, as required by the part 25 sections listed in paragraph 4 of this AC. Many of the conditions for which this AC provides guidance pertain only to certain 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 Nominal Conditions.

For nominal conditions without failures, malfunctions, or adverse conditions, the applicant must, per § 25.629, show freedom from aeroelastic instability 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 of this AC 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.

AC 25.629-1C

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 Maximum Mach Number.

The applicant may limit the aeroelastic stability envelope 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 Configurations.

Some configurations and conditions that § 25.629 and other part 25 regulations require to be investigated are 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 of this AC shows the minimum aeroelastic stability envelope for fail-safe conditions. This envelope 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.

AC 25.629-1C

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 + .05MC

6.2 Configurations and Conditions.

The following paragraphs summarize the configurations and conditions that applicants should investigate in demonstrating compliance with part 25. If aeroelastic stability is sensitive to load conditions of operation within the airplane’s maneuvering envelope, the applicant should show the airplane’s freedom from aeroelastic instability for any condition of operation allowed by the maneuvering envelope, using rational analyses, parameter variations, or an appropriate test. 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. The applicant should show the airplane’s freedom from aeroelastic instability throughout the expanded clearance envelope described in paragraph 6.1.1 of this AC 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.

AC 25.629-1C

6.2.1.3 All normal combinations of autopilot, yaw damper, or other automatic flight control systems.

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, to address the influence of gyroscopic loads and thrust on aeroelastic stability.

6.2.2 Sensitivity to Load Conditions.

Sensitivity of aeroelastic stability to load conditions can be determined by varying aerodynamic coefficients and conducting other parametric studies to assess the effect of redistribution of air loads due to structural and control surface deflections. Special configurations may require additional studies.

6.2.3 Failures, Malfunctions, and Adverse Conditions.

The applicant should investigate the following conditions for aeroelastic instability within the fail-safe envelope defined in paragraph 6.1.3 of this AC.

6.2.3.1 Any critical fuel loading conditions, not shown to be extremely improbable, which may result from mismanagement of fuel.

6.2.3.2 Any single failure in any flutter control system.

6.2.3.3 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.

6.2.3.4 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).

6.2.3.5 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.

6.2.3.6 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 of this AC.

6.2.3.7 Any single propeller or rotating device capable of significant dynamic forces rotating at the highest likely overspeed.

AC 25.629-1C

6.2.3.8 Any damage or failure condition required or selected for investigation by § 25.571. The applicant need not consider the single structural failures described in paragraphs 6.2.2.4 and 6.2.2.5 of this AC in showing compliance with this paragraph if—

6.2.3.8.1 The structural element could not fail due to discrete source damage resulting from the conditions described in § 25.571(e); and

6.2.3.8.2 A damage tolerance investigation in accordance with § 25.571(b) shows that the maximum extent of damage assumed for the purpose of residual strength evaluation does not involve complete failure of the structural element.

6.2.3.9 The following flight control system failure combinations where aeroelastic stability relies on flight control system stiffness, damping or both:

(i) any dual hydraulic system failure;

(ii) any dual electrical system failure; and

(iii) any single failure in combination with any probable hydraulic system or electrical system failure.

6.2.3.10 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.

6.2.3.11 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.

Applicants should design main surfaces, such as wings and stabilizers, to meet the aeroelastic stability criteria for nominal conditions, and investigate those surfaces 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.

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, the applicant should consider failures of those components possible adverse effects.

AC 25.629-1C

6.3.3 Reliance on Control System Stiffness.

Where aeroelastic stability relies on control system stiffness and/or damping, the applicant must consider additional conditions in accordance with § 25.629(d). This includes that—

6.3.3.1 The actuation system continuously provides, at least, the minimum stiffness or damping required for showing aeroelastic stability for the following conditions—

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,

25.671, and 25.1309.

6.3.3.2 The actuation system minimum requirements must also be continuously met after any combination of failures not shown to be extremely improbable (occurrence less than 10-9 per flight hour). As required by §§ 25.629(d) and 25.1309, applicants must conduct a qualitative assessment in addition to the quantitative assessment. Applicants must also consider the latent failure criteria of § 25.1309(b)(4) and (5). In accordance with § 25.629(d)(9), certain combinations of failures, such as dual electric or dual hydraulic system failures (including loss of hydraulic fluid), or any single failure in combination with any probable electric or hydraulic system failure (including loss of fluid), are assumed to occur regardless of probability calculations and must be evaluated. 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 Freeplay.

Applicants may incorporate consideration of freeplay 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 Balance Weights.

If a proposed design uses balance weights on control surfaces, the applicant should substantiate the effectiveness and strength of those weights and their support structure.

AC 25.629-1C

6.3.6 Automatic Flight Control System.

The automatic flight control system should not interact with the airframe to produce aeroelastic instability, in this case known as aeroservoelastic instability. To meet the requirements of § 25.629, in the nominal condition, the structural modes must have adequate aeroservoelastic gain and phase stability margins for any single control system feedback loop at speeds up to VD/MD, or for control laws with limited operating envelopes, the control law design speed may be used in place of VD/MD. When analyses indicate possible adverse coupling, the applicant should perform tests 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.

Applicants may demonstrate compliance with aeroelastic stability requirements for an airplane configuration by analyses, tests, or combination thereof. In most instances, analyses will be necessary to determine aeroelastic stability margins for normal operations, as well as for possible failure conditions. Applicants may use wind tunnel flutter model tests, where applicable, 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.

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 during 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.

AC 25.629-1C

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, as described below, consistent with configuration complexity, will be 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 GVTs, 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 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.