U.S. Department Advisory of Transportation Federal Aviation Circular Administration Subject: Aeroelastic Stability Substantiation of Date: 10/27/14 AC No: 25.629-1B Transport Category Airplanes 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 of this Michael Tossset Michael Kaszycki Acting Manager, Transport Airplane Directorate Aircraft Certification Service
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
6.2.1 Nominal Configurations and Conditions5
6.2.2 Failures, Malfunctions, and Adverse Conditions6
6.3 Detail Design Requirements7
7 Compliance8
7.1 Analytical Investigations8
7.1.1 Analytical Modeling9
7.1.2 Types of Analyses10
7.1.3 Damping Requirements10
7.1.4 Analysis Considerations12
7.2 Testing15
7.2.1 Structural Component Tests16
7.2.2 Control System Component Tests16
7.2.3 Ground Vibration Tests16
7.2.4 Flutter Model Tests17
7.2.5 Flight Flutter Tests18
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Contents (continued) 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 111
Figure 5. Damping versus Velocity—Method 212
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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.
• 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.
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.
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.
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.
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.
MD
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
1.15 VC VD + .05MC Airspeed
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.
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 6
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 7
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.
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.
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 9