80AFRC24R0008-Appendix 2_AFG-7123.1-001.pdf

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Armstrong Flight Research Center AFG-7123.1-001, Revision A-1 Edwards, California 93523 Expires October 1, 2025

SUBJECT:

Aircraft Structural Safety of Flight Guidelines

RESPONSIBLE OFFICE:

Research and Engineering Directorate

Aircraft Structural Safety of Flight Guidelines AFG-7123.1-001, Revision A-1 Expires October 1, 2025

CONTENTS

1.0 PURPOSE & OBJECTIVES

2.0 SCOPE & APPLICABILITY

2.1. Scope

2.2. Applicability

3.0 INTRODUCTION

4.0 SUMMARY

4.1. Static Structural Design

4.2. External Loads

4.3. Structural Dynamics and Aeroelasticity

4.4. Thermal Loads

5.0 STATIC STRUCTURAL DESIGN GUIDELINES

5.1. Design and Analysis

5.2. Ground Loads Testing

5.3. Instrumentation and Monitoring

5.4. Structural Inspection

5.5. Flight Envelope Limitation

5.6. Fatigue Life Considerations

5.7. Military Standard Comparison

5.8. COMPOSITES

5.8.1. Introduction

5.8.2. Design / Analysis

5.8.3. Design Allowables

5.8.4. Environmental Considerations

5.8.5. Damage Tolerance

5.8.6. Substantiation / building block approach (testing)

5.8.7. Manufacturing

5.8.8. Inspection / acceptance

5.9. UNIQUE MATERIALS

5.9.1. Introduction

5.9.2. Material Design Considerations

5.9.3. Manufacturability and Confidence

6.0 EXTERNAL LOADS

6.1. Overview

6.2. Introduction

6.3. External Loads Analysis

6.3.1. General

6.3.2. Types of Aircraft Load Cases

6.3.3. Flight Condition Selection

6.3.4. Load Analysis Applications

6.4. Loads Report Content

content. This document may be distributed outside of the Center.

6.5. References for External Loads Analysis

6.6. Summary

7.0 STRUCTURAL DYNAMICS AND AEROELASTICITY

7.1. Introduction

7.2. Summary of Aeroelastic Instabilities

7.3. Minimum Flutter Margins

7.4. Flutter Clearance Approaches

7.5. Flight Regimes

7.6. Ground Vibration Test

7.7. Control Surface Free-Play Test

7.8. Structural Mode Interaction Test

7.9. Finite Element Model Tuning

7.10. Flight Flutter Test

7.11. Documentation

8.0 THERMAL LOADS

8.1. Analysis Guidelines

8.2. Test Guidelines

8.3. Thermal Factors of Safety and Margins of Safety

9.0 RELEVANT DOCUMENTS

9.1. Authority Documents

9.2. Referenced Documents

9.3. Informational Documents

APPENDIX A, Definitions

APPENDIX B, Acronyms content. This document may be distributed outside of the Center.

1.0 PURPOSE & OBJECTIVES

The purpose of this document is to provide guidelines for the structural design of experimental aircraft, aircraft structural components, and structural modifications to existing experimental aircraft operated at Armstrong Flight Research Center (AFRC, hereinafter referred to as the Center), in Edwards, California. These guidelines account for structural instrumentation options, proof test options, and flight test operational envelope options, all of which affect the initial structural design approach. The guidelines in this document can be tailored based on the risk posture of an individual project.

2.0 SCOPE & APPLICABILITY

2.1. Scope

Modifications to National Aeronautics and Space Administration (NASA) Center aircraft to include installation of payloads, and modification or replacement of original aircraft structure, except when covered by technical orders, service and repair manuals, airworthiness directives, and other similar documents issued by original manufacturers of aircraft, certified under military or civilian airworthiness authorities. Scope also includes the design of new special purpose or research aircraft not certified under military or civilian airworthiness authorities.

2.2. Applicability

This guidance applies to all organizations engaged in the structural design, analysis, and modification (including payload installation) of Center aircraft including Aerostructures Branch, Operations Engineering Branch, and Programs and Projects Directorate, as well as customers, partners, and contractors supporting NASA and the Center in this work.

3.0 INTRODUCTION

The Center is engaged in flying many varied and unique aircraft. Frequently, these aircraft require structural modifications in order to accommodate unique flight experiments; and sometimes, a completely new one-of-a-kind experimental airplane is built. Whenever a new airplane, or one with substantial modifications such as new lifting surfaces or new control laws, is flight-tested, structural loads and dynamics flight-testing become an important element of the airworthiness approach.

There are five distinct objectives associated with the Center’s structural engineering:

A. Acquire stress, load, temperature, deflection, mode shape/frequency or damping data for verification of design predictions or analysis codes.

B. Safely expand the flight envelope of research aircraft to enable the acquisition of data required for performance, stability and control, propulsion, and any other disciplinary research objectives.

C. Demonstrate new structural materials and structural concepts in the flight environment.

D. Demonstrate that a prototype vehicle can perform all military standard or Federal Aviation Administration (FAA) regulation maneuvers everywhere in the design flight envelope. This process is often referred to as “structural demonstration.”

E. Support the Center process for self-certification of airworthiness of an airborne science platform aircraft and maintain its structural safety of flight over its lifetime as its on-board experiments or missions evolve.

The first two objectives are, historically, the most common types of flight activity at the Center. The Center has been a leader in the development of flight and ground load, deflection and vibration measurement techniques, and has developed an experienced staff for generating analytical predictions, planning load and aeroelastic buildup flights, monitoring loads and aeroelastic effects in real time, and conducting post flight data analysis in support of safety of flight and for flight research documentation and reporting. These activities often involve evaluating new and unique configurations such as forward swept wings, oblique wings, joined wings, advanced flexible wings, etc., which are too immature for commercial design applications when they are first committed to flight. Emphasis is placed on correlation with analytical and wind tunnel predictions and ground test results, and the development of new instrumentation and flight test techniques with a thorough understanding of the phenomena involved so that a dependable assessment of the generic technology can be developed.

The third objective occurs less frequently and generally involves little or no instrumentation. The materials and concepts are usually developed under contract to NASA or another government entity and the responsibility for justifying their flight worthiness rests with contractors. These activities usually involve the design, analysis, and ground test of panels or small components, such as control surfaces, which are subsequently mounted on existing vehicles and flight-tested.

The fourth objective is usually associated with satisfying a contractual procurement specification for a production aircraft. It is conducted by the airframe manufacturer using private or government facilities on an intensive flight test schedule. The airframe contractor assembles a special structural demonstration flight-test team consisting of highly experienced structural flight test engineers backed up by the structural designers and analysts who designed and developed the airplane. The goal of this activity is to content. This document may be distributed outside of the Center.

force a relatively mature configuration to the limits of its design envelope as soon as possible, thus demonstrating both the adequacy of the airplane structure and the predicted flight environment (aerodynamic flow field and airplane kinematic responses).

This is one of the highest risk types of flight testing.

The fifth objective often involves multiple structural engineering specialties. It may be required to develop new external loads, new internal loads, design a structure, and perhaps test it before it is installed. Flutter clearance or thermal analysis may be required.

4.0 SUMMARY

This summary will provide the highlights of the Center’s structural guidelines. The details of these guidelines will be further discussed in later sections.

4.1. Static Structural Design

All static structural designs require some factor of safety (FS) in order to lower the probability of in-flight structural failures. Different materials, structural analysis, and previous load testing are all part of determining the minimum FS a structure will require for flight. Proof load testing and using identical structural design to previously loaded structures can lower the FS that is necessary. Table 1 shows a summary of the static structural design guidelines for common materials and conditions. Sections 5.8 and

5.9 will discuss the use of composite materials and unique materials.

Unique materials can be used to meet specific mechanical property needs in a structure. The use of these unique materials will require specialized, project-oriented requirements and guidelines.

Table 1: Static Structural Design Summary

Factor of Safety Material Condition See Section

2.25 ~ 3.0 Composite Structure verified by analysis along with building block approach

5.1

2.25 Metal Structures verified by analysis only

1.80 Metal or

Composite

Structure verified by proof tests up to 120% Design Limit Load (DLL)

1.50 Metal or

Composite

Structural proof test plus full flight instrumentation additional 1.15

Joints and Fittings

Where failure of one fastener, pin, or lug could result in loss of a component content. This document may be distributed outside of the Center.

4.2. External Loads

An external loads analysis predicts the flight loads acting on an aircraft by the environment. The external loads analysis is necessary when there is any change made to the outer mold line (OLM) of an aircraft. This analysis will determine if the modification is structurally sound and will not adversely affect the baseline aircraft. The analysis calculates the aerodynamic loading and inertial loading of the aircraft based on worst-case flight conditions. Engineering judgement is necessary to select the minimum amount of analysis needed that will bound any loads experienced during flight. Section 6.4 identifies the contents of a loads report required for a complete airworthiness assessment.

4.3. Structural Dynamics and Aeroelasticity

Aircraft have the potential to exhibit aeroelastic and aeroservoelastic (ASE) instabilities which can be prevented by following the tables below.

Preventing these aeroelastic instabilities relies on thorough analysis and testing procedures. Table 2 shows the different airworthiness approaches which can be used for the flutter clearance of an aircraft based on the information available. Table 3 shows the essential aeroelasticity guidelines for performing flutter analyses and ASE analyses for flight clearance approval. Table 4 shows the guidelines for finite element model (FEM) updating or tuning which can be done using either the military standard (MIL-STD)-1540, Product Verification Requirements for Launch, Upper-stage, and Space Vehicles or the NASA standard (NASA-STD)- 5002, Load Analyses of Spacecraft and Payloads. Engineering judgement is necessary to tailor and select which airworthiness approach and model updating standard as appropriate for each project.

Table 2: Flutter Clearance Approaches

Airworthiness Approach See Section

Flutter Clearance

1. Aeroelastic similarity

7.4 2. Analytical parametric investigation

3. Flutter analysis, ground vibration testing (GVT), and flight flutter testing content. This document may be distributed outside of the Center.

Table 3: Aeroelasticity Guidelines Guidelines See Section

Flutter Clearance Margins Minimum flutter margin of 15% 7.3 Aeroservoelastic Stability Margins

Gain margin (GM) of at least ±6 decibel (dB) and at least ±45° phase margin

(PM) on each feedback loop for closed loop flutter if the frequency is higher than

0.06 Hertz, per second (Hz) and lower

than the 1st flutter frequency and the speed is higher than the minimum operating speed and lower than the maximum operating speed.

7.3

Table 2: Finite Element Model Tuning Guidelines

Guidelines See Section Finite Element Model Tuning:

Use engineering judgement to select which of the two standards is applicable

1. MIL-STD-1540,

Product Verification Requirements for Launch, Upper-stage, and Space Vehicles

a. Analytical model frequencies are to be within 3% of test frequencies.

7.9

b. Using a cross-orthogonality matrix formed from the analytical mass matrix and the analytical and test modes, corresponding modes are to exhibit at least 95% correlation and dissimilar modes are to be orthogonal to within 10%.

2. NASA-STD-5002,

NASA Technical Standard; Load Analysis of Spacecraft and Payloads

a. Agreement between test and analysis natural frequencies should, as a goal, be within 5% for significant modes.

7.9

b. Accurate mass representation of the test article should be demonstrated with orthogonality checks using the analytical mass matrix and the GVT test mode shapes. The orthogonality matrix off-diagonal terms should be less than 0.1 for significant modes based on the diagonal terms normalized to 1.0.

c. For mode shape comparisons the absolute value of the cross-orthogonality matrix between corresponding test and analytical mode shapes should be greater than content. This document may be distributed outside of the Center.

0.9; and all other terms of the matrix should be less than 0.1 for all significant modes.

4.4. Thermal Loads

Required factors of safety vary by application. The requirements will be identified in the Project’s Systems Requirement Document (SRD). See section 8.0 for more details.

5.0 STATIC STRUCTURAL DESIGN GUIDELINES

Design and analysis, static ground loads testing, instrumentation and monitoring, structural inspection, and flight envelope limitations, among others, are major relevant processes that can be implemented in appropriate combinations to enable a flight research project to safely achieve its flight goals. These guidelines have been applied to past and present Center aircraft flight programs and are planned for application to future programs.

5.1. Design and Analysis

It has been shown historically that if all other variables are held constant, the higher the structural design FS is, the lower the probability of in-flight structural failure will be. Because of the routine operating cost of increased structural weight and the effect of the diminishing return of large design factors, the aviation industry has sought out an acceptable or optimum value for the design FS. It should be noted that this chosen value is nearly always larger than that which is required to avoid material yielding at design limit load. This is due to considerations such as fatigue life, operational durability, and joint inefficiency. While designers of production aircraft normally expend considerable effort to minimize the aircraft structural weight fraction because of the importance of vehicle performance or operating efficiency, the designer of a research modification to an existing aircraft, or even a one-of-a kind research aircraft, can often tolerate considerable structural weight inefficiency.

External loads analysis is used to calculate the aggregate air load acting on the outer surface of a flight vehicle. Hand analysis using aerodynamic handbook tables can yield predicted approximate values in some cases.

Computational fluid dynamic (CFD) modeling is often used for complex or structurally sensitive geometries. Similarity to previously flight-tested aircraft is valuable.

Internal loads analysis takes the external loads and the other loads and calculates the individual structural component loads. In statically content. This document may be distributed outside of the Center.

determinant structures, this may be done by hand analysis using basic principles but will often involve finite element analysis.

Adequate reporting of the external and internal loads analyses is needed to document the work and support any independent review.

Stress analysis takes component dimensions and material properties, uses this information to calculate component strength, compares this to the component net load including FS and produces a predicted MS.

Stress analysis approach may, but not required, follow these references:

1. Niu, Michael C.Y., Airframe Stress Analysis and Sizing, 3rd edition, Adaso/Adastra Engineering Center, Granada Hills, CA. 2011

2. Bruhn E.F., Analysis and Design of Flight Vehicle Structures, 2nd edition Adaso/Adastra Engineering Center. Granada Hills, CA. 2011

Material properties should be taken from current version of MMPDS-06, Metallic Materials Properties Development and Standardization (formerly Military Handbook (MIL-HDBK)5) or CMH-17, Composite Materials Handbook (formerly MIL-HDBK-17) when available. Using manufacturer provided proven load limits when available is preferred. For composite materials and unique materials, see sections 5.8 and 5.9 for details. The stress analysis report may include handwritten work and should present a clear table of the minimum margins found. MS is defined as: MS = (allowable load / ultimate load) - 1. Ultimate load is defined as: ultimate load = FS x design limit load. Allowable load is defined as the maximum load that a structure or structural component can sustain before resulting in a failed condition. Design limit load is defined as the worst possible loading conditions the aircraft structure and/or structural component is expected to encounter while operating within its allowable flight envelope.

The stress report should include enough information so that an independent check could be done. This provides some confidence in the strength of a structural article, but in more complex designs a load test of a physical article is sometimes needed. If a middle ground solution is appropriate, overall stress analysis combined with the test to failure of a critical sub- component (sacrificial article) may suffice to augment confidence in predicted strength. Note that the results of a stress analysis are never any better than the confidence available in the load values used in that analysis.

Care should be taken to explicitly recognize stiffness design requirements as well as strength design requirements when this is an issue.

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Batch material property certification from the vendor should be specified as a requirement during purchase. Load carrying parts should be of aircraft grade structural materials.

Fasteners intended for aircraft use should be specified to have 160 KSI (minimum) strength or greater as appropriate and should include fastener retention features. Over time as panels and assemblies, etc., are removed and replaced, it is possible for fasteners to migrate to other locations, therefore it is important that they all meet the minimum strength guideline. Without retention features fasteners will risk eventually becoming foreign object damage (FOD). It is recommended that corrosion resistant fasteners are selected, and that compatibility with adjacent materials be considered to avoid problems like galvanic corrosion.

A concise summary of recommendations for primary structure design factors of safety is given here:

A. 2.25 FS – for metallic structures if structural design is verified by analysis only.

B. 2.25 to 3.00 FS – for composite structures if structural design is verified by analysis along with building block approach for the design and analysis. For example, 2.25 FS can be used when using well established composite processes and materials, plus the full utilization of building block approach (section 5.8.3). Higher FS should be used when a limited building block approach is employed.

C. 1.80 FS – for either metallic or composite structure (when using well established composite processes and materials) when verified by proof tests to 120% of flight limit loads.

D. 1.50 FS – for either metallic or composite structures (when using well established composite processes and materials) when an identical test article has been loaded to 100% DLL in each of the various design cases with no yielding and when subsequently loaded without failure to 150% DLL using the most critical load case. The flight airframe should then be instrumented for loads monitoring during envelope expansion (Ref MIL-STD-8860).

E. An additional joint or fitting factor of not less than 1.15 FS should be used on all joints and fittings where failure of one fastener, pin, or lug could result in loss of a component or any major portion thereof.

Primary structure is any detail, element, or assembly that contributes significantly to the carrying of flight, ground, pressure, or control loads, and whose failure could affect the structural integrity necessary for the safety of the aircraft. Failure of primary structure compromises a vehicle’s structural integrity and therefore the ability to fly and land safely.

Secondary structure is a structure that is not a primary load-carrying member. Failure of secondary structure neither reduces the airframe structural integrity, nor prevents the airplane from continued safe flight and landing.

All primary structures should be designed for worst-case loads conditions, including an additional appropriate margin for uncertainty in the loads case. All secondary structures, the failure of which could result in damage to the primary structure, should be designed to the same specifications as primary structures. The design loads cases are subject to review and approval by appropriate Center structures engineers.

Beyond showing a non-negative margin on ultimate load, it should be shown that structural yielding will not occur within the design limit load.

Just as ultimate load is checked against ultimate strength, limit load should be checked against yield strength.

A proper stress analysis report should include:

1. A statement of the source of the external loads used

2. An explanation of the derivation of the internal loads used

3. A statement of the source of the material properties used

4. A statement of the factors of safety and any fitting factors used

5. An explicit statement of any assumptions used

6. References for any complex or empirical stress equations used

7. Comments or calculations on any fatigue concerns

8. A summary of the smallest margins of safety calculated

The margin summary should state the load case that caused the stress and the failure mode (tension, buckling, bending or shear).

5.2. Static Ground Loads Testing

A common objective of ground loads testing is the establishment of confidence in strength of a test structure. A proof test to failure reveals both design and fabrication flaws and material property deficiencies that are hard to find through other processes. Ground load testing can be used to augment stress analysis. The highest confidence in determining structural strength is provided by a proof test to failure of an equivalent test article. The results of a loads test are never any better than the confidence available in the load values chosen to be applied in that test.

content. This document may be distributed outside of the Center.

Both ultimate strength and failure modes can be confirmed by this activity.

A typical proof test load level is 150% of design limit load.

When the cost of failure testing a sacrificial test article is deemed too high, the flight article itself is sometimes tested to load levels that are predicted by analysis to be less than yield load. Typical peak load levels for qualification testing are on the order of 120% of design limit. While this level of load testing does not prove ultimate strength directly or show failure modes, it does add valuable confidence in a working load limit. It also adds confidence in the strength of a given flight article in the case where there may be concerns regarding fabrication process quality control from one unit or airframe to the next. This is especially useful when new material systems are being employed. Note that multiple load tests may be called for to establish confidence in strength of critical wing, fuselage, empennage, or landing gear load cases. If adequate local strain sensors are installed on the ground test article the FEM may be correlated to the test data produced. It is common to measure deflections of the test article under load to provide measured stiffness data as well.

Another objective of ground load testing may be the calibration of strain sensors by the application of known loads.

It should be noted that telemetry data system channel calibration is not the same as strain gage load calibration.

Any type of load testing should include appropriate structural inspection before and after the test.

5.3. Instrumentation and Monitoring

The highest confidence in determining aircraft design limit loads is established by a flight loads survey performed by an instrumented and load calibrated aircraft where the component loads are actually measured at the specific flight conditions. This can include uncalibrated local strain measurements as well as strain sensors calibrated to provide load indications in pounds or inch-pounds, etc.

The usual focus of structural instrumentation will be primary structure load paths especially at critical stations such as those where the manufacturer has specified load limits or in new structures where stress analysis has shown the smallest margins. In the case of some flight experiments, it makes sense to design the load path of a test structure to facilitate the needed load measurement.

content. This document may be distributed outside of the Center.

Sensor selection usually includes metallic foil strain gages, but optical fiber-based systems are now sometimes considered for specific benefits.

Measured surface air pressures can sometimes be integrated and used to calculate external loads. Thermocouples are used when material temperatures or thermal stresses are an issue. Where metallic strain gages are considered, full four active arm bridges are preferred for sensitivity and to reduce thermally-induced errors. If a single or two active arm bridge configuration is used, local bridge completion with the passive gages strain isolated but thermally in contact with the structure is strongly preferred. Where the local principle axis is not well known, three arm rosettes can be employed. This is a configuration with three single active arm bridges with the active legs oriented at 0, 45 and 90 degrees with the rest of each bridge being locally completed as just discussed. Strain instrumentation can be used to better understand local load paths.

When monitoring real-time flight loads or strains it is often helpful to display these values as percent of limits. When monitoring loads, it is very valuable to have the manufacturer’s established load limits if they are available. Control room displays should be designed to provide moment by moment awareness of the stress state of the flight structure so that the aircraft can be protected from load limit exceedance. Another valuable display feature is a “hold last peak value” capability so that even a momentary load peak or a limit exceedance will not be missed in real time.

A ten second dwell time is often used with this feature. In this manner, both current load values and recent peaks are shown in relation to corresponding strength limits. Of course, a primary goal of real time monitoring is to perceive and track load trends so that load limit exceedances can be precluded.

Post flight analysis is often used to examine more complex relationships and to prepare flight reports or to perform research analysis.

5.4. Structural Inspection

This pre and /or post-flight process is intended to help verify structural soundness or to detect structural damage before it can pose a threat to human life or expensive resources. While some damage is abrupt and catastrophic in nature, structural failure is often progressive, beginning with a small but detectable crack or damage to a single fastener or subcomponent. Techniques include: Visual, Dye Penetrant, Magnetic Particle, X-Ray, Ultrasonic, Eddy Current, Thermography. Production aircraft come from the manufacturer with an established inspection plan.

When such an aircraft is modified, or its operational employment deviates from its intended routine use, it is appropriate to modify the inspection plan. New structures or modifications should be designed to be content. This document may be distributed outside of the Center.

inspectable. The inspection plan should be briefed at appropriate levels of detail at preliminary design review (PDR), critical design review (CDR) and flight readiness review (FRR). Inspection findings should be communicated to the Flight Operations Engineer and the Structures Engineer and briefed at Tech Briefs and Crew Briefs.

5.5. Flight Envelope Limitation

It is sometimes easier to reduce load than it is to increase strength if a negative margin is found. Sometimes this is done by reducing the allowable normal acceleration (Nz) or the aircraft gross weight. In some cases, it is appropriate to reduce the allowable dynamic pressure or some other operational parameter. If this is done, the new limit should be clearly stated in the pilot’s handbook, etc., and briefed at the Tech Brief, Crew Briefs, and so on. Sometimes it is helpful to note these limits on the flight cards.

In some cases this limit may be progressively altered by point-to-point clearances during flight or in a flight-to-flight manner when appropriate measurements may be compared against known limits.

5.6. Fatigue Life Considerations

Historically, the Center flight research projects have been very diverse, however, most have had relatively short lives with regard to total flight cycles or flight hours. In these cases, a project structures safety of flight plan that includes only a rudimentary fatigue analysis combined with periodic structural inspections has proven adequate to preclude fatigue failure problems. The analyses in these cases often show the structure good for a recommended scatter factor of four times the planned number of flight cycles or flight hours using a simple predicted load spectrum.

Occasionally however, a project comes along that requires a very long operational life involving thousands or tens of thousands of flight cycles with extensive new or modified structure and/or new load spectrum. In these cases, a much more elaborate and expensive effort is performed that involves instrumenting the aircraft for loads and performing flight loads surveys to measure load spectrums for all mission phases. The family of load spectrums is then used in an extensive fatigue and damage tolerance study. This study is then used to predict structural life, guide the creation of the new structural inspection plan and recommend critical areas for further modification for fatigue life extension if needed.

An example of unintended consequences with fatigue life is the case where a centerline pod is added beneath the fuselage upstream of an content. This document may be distributed outside of the Center.

existing blade antenna. The plan for the pod itself is well devised and executed but the effect of the pod’s wake on the antenna is overlooked.

With the new oscillatory air loads on the antenna it fatigues quickly and within about a dozen flights the antenna base or the adjacent fuselage skin cracks. The lesson learned here is that anything that changes the structural strength OR the load magnitude, distribution, or cyclic nature needs to be taken care of.

5.7. Military Standard Comparison

The integrated design criteria, proof-test and instrumentation guidelines, and corresponding operational flight envelopes are graphically illustrated in Figure 1.

Figure 1: MIL-A-8860, General Specifications for Airplane Strength and Rigidity, design criteria, instrumentation, proof test, and flight test guidelines for experimental aircraft (metallic only)

A. The first section of Figure 1 defines the approach needed to satisfy the MIL-A-8860 for the design, testing, and qualification of military content. This document may be distributed outside of the Center.

production aircraft. Typically, an FS of 1.5 is used, i.e., Ultimate Load = 1.5 x Limit Load. A positive, or zero, MS is maintained based on analysis.

The aircraft is fully instrumented for loads and stresses, i.e., strain gages are installed on all lifting surfaces, control surfaces, and the fuselage. The strain gages are calibrated for lifting surface loads, control surface hinge moments, and fuselage loads. These strain gages are calibrated by the application of known loads and the derivation of load equations. Additional strain gages are installed for local stress measurements. A flight loads survey to 80% of limit is performed to produce high confidence flight loads using the measured flight data.

A dedicated static-test aircraft is fully instrumented (about 10 times the instrumentation of the flight-test aircraft) and ground load tested to ultimate load. The flight-test aircraft is then used to perform flight structural demonstration tests to limit load to certify the aircraft. Once certified, all aircraft incorporating the same basic structural design and flight control laws are considered cleared to operate within the certified envelope without further testing.

B. The second section of figure 1 shows option 1. In this case, a design FS of 2.25 is used and a non-negative MS is maintained.

Conservative load predictions are used. No strain gage/loads instrumentation is installed and no proof tests are performed. The aircraft is operated such that none of the predicted design limit loads are exceeded. An example similar to this option is the AFTI/F-111 Mission Adaptive Wing leading edge and trailing edge devices.

C. The third section of figure 1 shows option 2. Typically, an FS of 1.5 is used, and a non-negative MS is maintained. The aircraft is fully instrumented with calibrated strain gages, and the flight aircraft is qualification ground load tested to design limit load. The aircraft is cleared to fly to 80% of the tested load through a nominal envelope expansion process. Clearance may be given to fly to 100% limit load on a case-by-case basis. An example of this option is the F-8 Supercritical Wing Research Aircraft.

D. The fourth section of figure 1 shows option 3. In this case a design FS of 1.8 is used and a non-negative MS is maintained. Conservative load predictions are used. With conservative loads predictions, no strain gage/loads instrumentation is required. However, many projects use calibrated loads instrumentation since there is often low confidence in the loads predictions. The flight article is qualification ground load tested to 1.2 DLL and inspected to verify that no yielding or other damage has occurred. The aircraft is operated inside the content. This document may be distributed outside of the Center.

analyzed operating envelope such that none of the predicted design limit loads are exceeded.

E. The fifth section of figure 1 shows option 4. For this option, the project proposes its approach to demonstrate structural airworthiness, which includes but not limited to FS, proof test, instrumentation, and loads predictions. It is strongly recommended that projects reach out to the Center Aerostructures Branch early in the program to discuss airworthiness approach to guide design, analysis, and testing efforts.

In all options appropriate structural inspection plans are created and executed.

Option 1 is generally used when modifying an aircraft that has already met Military Standard or FAA Regulation flight qualification requirements, and the modifications are simple with easily determined load paths, the design load predictions are easily derived or are known to be conservative, and component proof tests may have been conducted to verify the design concept in the laboratory.

When option 1 is used it is usually because typical structural modifications to existing aircraft using the conservative design criteria do not result in a significant weight penalty but does result in significant cost savings in avoiding instrumentation installations and proof tests. This option also provides a reasonable flight operational envelope. A variation on this combination could include the use of a smaller design FS with a corresponding reduced operating envelope such that the ratio of the FS used to the percent of design limit load flown to remains 2.25 or greater.

An example of this would be designing to 1.5 but operating to only 60% of limit. Of course, option 1 (and option 3 as well) is only safe to use if conservative load predictions can be made with confidence. If the loads are uncertain then appropriate instrumentation and monitoring should be used.

Not all aircraft designs or modifications can fall into option 1. Most of the one-of-a-kind research aircraft flown at the Center have used option 2 criteria.

The goal of applying these guideline combinations is to put together a project-specific structures safety of flight plan that accomplishes the project’s goals at minimum cost and effort while providing adequate confidence in structural strength and load. To this end, it is recommended that new or proposed flight projects request direct review or consultation from the Aerostructures Branch early in their planning process so that they content. This document may be distributed outside of the Center.

will have good prospects for timely approval to fly by the Center processes and a safe flight campaign at minimum cost and effort.

5.8. COMPOSITES

5.8.1. Introduction

Manufacturers are taking more advantage of composites due to their high strength to weight ratio, corrosion resistance, low conductivity, and nearly perfect elastic properties. However, composite materials and structures differ significantly from their metallic counterparts. Examples of such differences include variability in material properties, damage tolerance, additional failure modes, environmental effects, and the need for process control during manufacturing. Due to these differences, composite materials and structures are treated and looked at differently than metallics. Additional guidance on the quality system for the manufacture of composite structures can be found in FAA Advisory circular (AC) 21-26A, Quality Control for the Manufacture of Composite Structures.

5.8.2. Design / Analysis

The purpose of this analysis subsection is not to prescribe how to design or analyze specific parts, but rather what material properties and criteria should be used in those efforts. Validation of the design and analysis is done by utilizing a building block approach which is covered in a later section.

5.8.3. Design Allowables

As previously mentioned, composite material properties exhibit a large amount of variability which requires database design values to be determined statistically from test data. For reference, CMH- 17 pedigree requires 30 coupons per condition per material property for B-Basis values. Due to the large costs associated with generating true database design allowables, reduced data sampling and statistical extrapolations can be made to determine said design allowables. This, however, is reviewed on a case by case basis. It is advisable to use elevated temperature wet material property values due to environmental conditions at Edwards Air Force Base (EAFB, hereinafter referred to as Air Force Test Center (AFTC)), which is addressed in a later section.

A concise summary of design allowables are given here:

• Primary structure – A-Basis o 99% probability with 95% confidence interval

• Secondary structure – B-Basis o 90% probability with 95% confidence interval

5.8.4. Environmental Considerations

Weather conditions at AFTC are typically hot and arid, though the Center often has aircraft on global deployments in different atmospheric conditions. While typically a very dry climate, AFTC is home to the occasional rainstorm which causes humidity, thus detrimentally affecting composite material properties. During the summer, temperatures can reach nearly 200°F on dark surfaces exposed to the sun. As such, it is imperative for the composite matrix to have a glass transition temperature well above 200°F to prevent softening or degraded properties of the composite laminate structure, or special operational precautions to limit ramp time and structural temperatures. As a general practice it is recommended that all composite surfaces be painted white to reduce surface temperatures.

Aircraft structures at AFTC are also subjected to intense ultra violet (UV) radiation from direct sunlight. Long term exposure results in damage to the composite matrix, and therefore any composite parts that are exposed to long term sunlight should be painted with a UV blocking paint.

5.8.5. Damage Tolerance

When composite structures are impacted with enough kinetic energy to cause damage, they do not exhibit damage like an aluminum structure, i.e., in the form of a visible dent. Especially in the case of painted parts, composite structures, when impacted may not show visible damage. However, the laminate or sandwich core can be delaminated or contain discontinuous fibers. The result of this is a structure with significantly altered material properties. It is therefore very important that composite structures be designed to be resistant to damage that naturally occurs with use. This design criterion is very application- and project-specific and needs to be evaluated on a case by case basis. Consult with the Aerostructures Branch for guidance regarding this subject.

5.8.6. Substantiation / building block approach (testing)

As previously mentioned, composite structures exhibit a large variety of failure modes and, depending on complexity, cannot be easily cleared through analysis alone without resorting to higher design FS. This is due to the inability of current finite element analyses to predict failure modes on large, complex structures. As a result, the design of composite structures needs to be substantiated through the “building block approach.”

This approach, which is graphically represented in Figure 2, employs testing of certain design features at each of the complexity levels which is then used to validate the design (or warrant a design change). In addition to validation/testing, this approach further reduces risk by ensuring the “manufacturability” of the design and allows the exercise of the materials and processes.

Figure 2: Building block approach pyramid of tests (ref. CMH-17)

As the Center is a research and development organization, the use of the building block approach is not intended to be cost prohibitive.

Testing should be kept to a minimum, yet sufficient level to support the design for its specific application and mission. Due to the uniqueness of each design and project, the building block approach needs to be evaluated on a case by case basis. It is imperative to consult with the Aerostructures Branch early in the development stages to establish project testing requirements in order to prevent costly and unnecessary project delays.

5.8.7. Manufacturing

With composites, the parts manufacturer is also a material manufacturer. This is in stark contrast to metal fabrication in which the part manufacturer fabricates parts from material supplied by the material supplier. The process in the latter scenario has little impact or well-defined changes to material properties. When making composite parts and structures, quality is extremely dependent on process and process variables. The issue is compounded by the fact that after a composite part is made, it is often difficult to inspect and assess it. It is critical that composite parts and structures are manufactured with strict adherence to process and process control. Therefore, material and process specifications used to produce composite structures should contain sufficient information to ensure that critical parameters in the manufacturing process are identified to facilitate production and final inspection. To ensure the quality, process procedures should clearly define specific materials, tooling, equipment, cure cycle parameters, quality standards, operator qualifications, storage and handling requirements, traceability records, and any other special requirements.

When manufacturing composite flight hardware, it is highly recommended to make witness coupons which can be tested, thus allowing direct insight into the quality of the parts.

Composite parts are to be built with material systems that have not expired, or exceeded out-life time limits, and are traceable to a manufacturer’s lot or batch number. Much like flight hardware that is manufactured from aluminum and requires a material certification, alloy, and temper, so too, composites require specific fiber type and weave architecture, resin type, ply schedule, and cure cycle.

5.8.8. Inspection / acceptance

Finally, it is highly recommended to develop an inspection plan for the acceptance or rejection of manufacturing-induced defects and damage, and the life of the composite structure. The plan should take into account the process and inspection capability. Visual inspections at times may be appropriate, given the nature of the structure, however, other non-destructive evaluation (NDE) methods may be required. When possible, one should establish a baseline on the composite structure immediately following the fabrication process. It is encouraged to make NDE “standards” with content. This document may be distributed outside of the Center.

known defects so when anomalies are found in the composite structure they can be quantified.

5.9. UNIQUE MATERIALS

5.9.1. Introduction

More advanced, highly specialized materials are being developed at a rapid pace. These materials can provide very distinctive mechanical properties and are being used in a wide variety of applications. However, these materials differ from metallic components. Each material has its own unique set of material properties and these properties should be well understood for the application of the material. Unique materials include, but are not limited to, advanced ceramics and additive manufactured materials.

5.9.2. Material Design Considerations

5.9.2.1. Advanced Ceramics

Advanced ceramics provide a lightweight, high melting temperature alternative to metallic components. These ceramics have high machinability, are chemically inert, good thermal and electrical conductors, and dimensional stability.

However, the production of such ceramics can be very complex and expensive. Prior analysis may not be available for certain material. Ceramics can also undergo static fatigue, which requires more inspections of the material.

Confidence in the manufacturing process and strength of the material is established to produce reliable parts.

5.9.2.2. Additive Manufactured Materials

Additive manufacturing, commonly referred to as 3- Dimensional (3D) printing, offers the ability to create highly complex geometries that traditional machining methods are unable to produce. There are various methods used to additively manufacture parts. Each method presents different challenges that are addressed when producing and testing parts. The repeatability of a printer is the key to producing reliable parts. The printer operator’s experience, type of printer, type of material, quality of material, printing process, and post processing treatments are all vital requirements for producing high quality parts. Much like composite materials, test coupons are manufactured alongside any potential part content. This document may be distributed outside of the Center.

and material properties tests are performed to identify any weaknesses.

5.9.3. Manufacturability and Confidence

Similar to composite materials, manufacturers making parts out of unique materials will, in most cases, also be the material manufacturers as well. The quality of the material manufacturing will determine the degree of confidence in the material. In the case of additive manufacturing, companies currently use material that is uncertified and changes depending on the printer, operator, and conditions during a print. This means that the material properties of a part can change from one print to the next, which creates an uncertainty in the reliability of parts. To minimize the associated risk of this uncertainty, the quality of the manufactured part is determined through testing before parts can confidently be used for any application. Additional inspections during the fabrication process and throughout the lifecycle of the part are highly recommended. Strict documentation and quality control procedures are agreed upon before the use of any unique material.

6.0 EXTERNAL LOADS

6.1. Overview

The Center specializes in flight research experiments. These research projects can range from the placement of a small pylon on an aircraft window frame, to a modified wing flap, all the way to new experimental aircraft. External loads analysis is a necessary step for many research projects in order to show safety standards are met. Whenever there is a modification of any size to the outer-mold-line of an aircraft, external loads analysis is performed to determine if the modification is structurally secure. Additionally, the analysis needs to be done to ensure that the modification will not damage the aircraft’s structure in flight. This section focuses on the background of external loads analysis, and how it is applied to aid research experiments at the Center.

6.2. Introduction

The Center has a need for external loads analysis in order to clear a modified aircraft for flight. The modifications can include a change in high lift device on the wing, a pod placed under a wing, a science…