513.8-22 tailoring guidance

MIL-STD-810H

METHOD 513.8

2.1.1 Effects of Acceleration

Acceleration results in loads on mounting hardware and internal loads within materiel. Note that all elements of the materiel are loaded, including fluids. The following is a partial list of detrimental effects from high levels of acceleration. If there is expectation that any of these may occur, it confirms the need to test for this effect.

• a. Structural deflections that interfere with materiel operation.
• b. Permanent deformation, structural cracks, and fractures that disable or destroy materiel.
• c. Broken fasteners and supports that result in loose parts within materiel.
• d. Broken mounting hardware that results in loose materiel within a platform.
• e. Electronic circuit boards that short out and circuits that open up.
• f. Inductances and capacitances that change value.
• g. Relays that open or close.
• h. Actuators and other mechanisms that bind.
• i. Seals that leak.
• j. Pressure and flow regulators that change value.
• k. Pumps that cavitate.
• l. Spools in servo valves that are displaced causing erratic and dangerous control system response.

2.1.2 Sequence Among Other Methods

• a. General. See Part One, paragraph 5.5.

b. Unique to this Method. Examine the life cycle environmental profile to determine the test sequence. Normally, acceleration is experienced after logistic storage and transportation environments and often near the end of the life cycle. Shock, vibration, and thermal stressing prior to acceleration testing is recommended as this will uncover failures that would not occur with unstressed items.

2.2.1 Procedure Selection Considerations

Subject materiel to be tested to both Procedures I and II tests unless otherwise specified. Subject manned aircraft materiel that is located in occupied areas or in egress and ingress routes to Procedure III. Procedure IV is special case for strength testing materiel that cannot be assessed by Procedures I and II.

2.2.2 Difference Among Procedures

2.2.2.1 Procedure I - Structural Test

Procedure I is used to demonstrate that materiel will structurally withstand the loads induced by in-service accelerations.

2.2.2.2 Procedure II - Operational Test

Procedure II is used to demonstrate that materiel will operate without degradation during and after being subjected to loads induced by in-service acceleration. Rationale for operational test levels can be found in paragaraph 6.1, reference k.

MIL-STD-810H

METHOD 513.8

2.2.2.3 Procedure III - Crash Hazard Acceleration Test

Procedure III is used to disclose structural failures of materiel that may present a hazard to personnel during or after a crash. This test is intended to verify that materiel mounting and/or restraining devices will not fail and that sub- elements are not ejected during a crash. Use for materiel mounted in flight occupied areas and/or that could block aircrew/passenger egress or rescue personnel ingress after a crash. The crash hazard can be evaluated by a static acceleration test (Method 513.8, Procedure III) and/or transient shock (Method 516.8, Procedure V). The requirement for one or both procedures must be evaluated based on the test item.

Only when the system and/or attachment method has a natural frequency below the knee frequency of the shock SRS, might this test be required to supplement the Crash Hazard Shock Test (see Method 516.8, Figure 516.8-9). For planning purposes, Procedure III should be included for budgeting and scheduling consideration until it is shown by analysis or a laboratory test that this procedure isn’t required.

2.2.2.4 Procedure IV – Strength Test

Procedure IV is a strength test primarily intended to generate specific loads in primary structures using sine burst testing. This method may be used as an alternative to static pull or centrifuge testing and is suitable for testing relatively stiff components, electronics boxes, instruments, or space vehicles. The sine burst test is a base-driven test in which the test article is subjected to a few cycles (typically 2 to 10 at peak amplitude) of sinusoidal input. This is usually done below the first resonant frequency of the test article to expose the hardware to a quasi-static loading. An example test profile is shown in Figure 513.8-1.

2.3.1 Test Axes

For the purpose of these tests, the axes should be consistent with the sign convention and axes used in the structural analysis of the platform with the direction of forward acceleration of the platform. The test item is tested in each direction along three mutually perpendicular axes for each test procedure. One axis is aligned with the forward acceleration of the platform (fore and aft, X), one axis is aligned with the span-wise direction of the platform (lateral, Y), and the third axis is perpendicular to the plane of the other two axes (up and down, Z). Positive rotational axes and accelerations vary between platforms as they are typically determined by various means such as use of the “left hand” or “right hand rule.” Figure 513.8-2 shows a typical vehicle acceleration axes system with sign convention defined by the “right hand rule”.

2.3.2 Test Levels and Conditions - General

Tables 513.8-I, 513.8-II, and 513.8-III & IV list test levels for Procedure I (Structural Test), Procedure II (Operational Test), and Procedure III (Crash Hazard Acceleration Test), respectively. When the orientation of the materiel item relative to the operational platform is unknown, the highest pertinent level applies to all test axes.

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MIL-STD-810H

METHOD 513.8

1/ Use levels specified for individual platforms and locations on/in the platforms. Use the values of this table only if platform criteria are unavailable.

2 Use levels in this column when forward acceleration is unknown. When the forward acceleration of the vehicle is known, use that value for A.

³ For carrier-based aircraft, use 4 as a minimum value for A, representing a basic condition associated with catapult launches.

4/ For attack and fighter aircraft, add pitch, yaw and roll accelerations as applicable (see paragraph 2.3.3).

For helicopters, forward acceleration is unrelated to acceleration in other directions. Test levels are based on current and near future helicopter design requirements.

6/ Manned vehicles capable of flight in the Earth's atmosphere and surrounding space environment.

When forward acceleration is not known, use the high value of the acceleration range.

8/ A is derived from the propulsion thrust curve data for maximum firing temperature.

2/ In some cases, the maximum maneuver acceleration and the maximum longitudinal acceleration will occur at the same time. When this occurs, test the materiel with the appropriate factors using the orientation and levels for the maximum (vertical) acceleration.

${ }^{10}$ Where $A^{\prime}$ is the maximum maneuver acceleration.

MIL-STD-810H

METHOD 513.8

1/ Use levels specified for individual platforms and locations on/in the platforms. Use the values of this table only if platform criteria are unavailable.

2 Use levels in this column when forward acceleration is unknown. When the forward acceleration of the vehicle is known, use that value for A.

³ For carrier-based aircraft, use 4 as a minimum value for A, representing a basic condition associated with catapult launches.

4/ For attack and fighter aircraft, add pitch, yaw and roll accelerations as applicable (see paragraph 2.3.3).

For helicopters, forward acceleration is unrelated to acceleration in other directions. Test levels are based on current and near future helicopter design requirements.

6 Manned vehicles capable of flight in the Earth's atmosphere and surrounding space environment.

7/ When forward acceleration is not known, use the high value of the acceleration range.

8/ A is derived from the propulsion thrust curve data for maximum firing temperature.

2' In some cases, the maximum maneuver acceleration and the maximum longitudinal acceleration will occur at the same time. When this occurs, test the materiel with the appropriate factors using the orientation and levels for the maximum (vertical) acceleration.

10/ Where A' is the maximum maneuver acceleration.

1/ Use levels specified for individual platforms and locations on/in the platforms. Use the values of this table only if platform criteria are unavailable.

2 The intent of this test is to disclose structural failures of materiel that may present a hazard to personnel during or after a crash. This test is intended to verify that materiel mounting and/or restraining devices will not fail and that sub-elements are not ejected during a crash. Use for materiel mounted in flight occupied areas and/or that could block aircrew/passenger egress or rescue personnel ingress after a crash.

3. Test item function is not required following this test. Thus test items that are not suitable for other tests or field use may be used for this test. Ensure test items are structurally representative (strength, stiffness, mass, and inertia) of the production design, but need not be functional. All contents (including fluids) designed to be carried in/on the materiel should be included. Tactical restraint systems should be loaded with the maximum expected weight.

See paragraph 6.1, reference a. Some system specific crash hazard acceleration levels are provided in Table 513.8-IV.

MIL-STD-810H

METHOD 513.8

OH-58D Kiowa Warrior Aircraft Crash Test Levels (g's) (Applied Separately) ²/

CH-47D/F Aircraft Crash Test Levels (g's) $\frac{3}{1}$

UH-1H Aircraft Crash Test Levels (g's)

UH/HH/MH-60A/L/M Aircraft Crash Test Levels (g's)

Figure description:

7 cycles, 22.5 Hz @ 10.08 Gpk

7 cycles, 22.5 Hz @ 10.08 Gpk

Figure 513.8-1. Example sine burst test specification for strength testing (from NASA-HDBK-7008).

Figure description:

The image illustrates the standard coordinate system and rotational moments for an aircraft, centered on its Center of Gravity.

Linear Axes:

• X-Axis (Longitudinal: +X is Fore, -X is Aft.
• Y-Axis (Lateral: +Y is Right, -Y is Left.
• Z-Axis (Vertical: +Z is Down, -Z is Up.

Rotational Moments:

• Rolling Moment (Φ: Rotation around the X-axis.
• Pitching Moment (θ: Rotation around the Y-axis.
• Yawing Moment (Ψ: Rotation around the Z-axis.

Figure 513.8-2. Typical directions of vehicle acceleration (right hand rule).

2.3.3 Test Levels and Conditions - Fighter and Attack Aircraft

The test levels as determined from Tables 513.8-I and 513.8-II are based on accelerations at the center of gravity (CG) of the platform. For fighter and attack aircraft, the test levels, must be increased for materiel that is located away from the vehicle CG to account for loads induced by roll, pitch, and yaw during maneuvers. When criteria are developed for specific aircraft, maneuver cases are considered and the resulting additional angular accelerations may add or subtract effects from the linear acceleration effects. When the following relationships (a-f) are used, it must be

assumed that the load factors always add. Thus absolute values are used in the equations. Add the load factors derived below to the Operational Test (Procedure II) levels of Table 513.8-II. Multiply the load factors derived below by 1.5 and add to the Structural Test (Procedure I) levels of Table 513.8-I. Do not add these values to the Crash Hazard Acceleration Test (Procedure III) levels of Tables 513.8-III and 513.8-IV.

a. Roll maneuver, up and down test direction. The additional load factor ($\Delta N_{Z}$) induced by roll, is computed as follows: $\Delta N_{Z} = (z/g) (d\ \phi / d\ t)^{2} + (y/g) d^{2} \phi / d\ t^{2}$

b. Roll maneuver, lateral left and lateral right directions. The additional load factor ($\Delta$N$_{\text{Y}}$) induced by roll, is computed as follows: $\Delta$N$_{\text{Y}}$ = (y/g) ($d$ $\phi$/$d$ t)$^{2}$ + (z/g) $d$ ^{2}$$\phi/$d$ t$^{2}$ - c. Pitch maneuver, up and down test directions. The additional load factor (ΔNz) induced by pitch change, is computed as follows: $\Delta \mathrm{N}_{\mathrm{z}}=(\mathrm{z} / \mathrm{g})(d \theta / d \mathrm{t})^{2}+(\mathrm{x} / \mathrm{g}) d^{2} \theta / d \mathrm{t}^{2}$ - d. Pitch maneuver, fore and aft test directions. The additional load factor (ΔNx) induced by pitch change, is computed as follows:$\Delta N_{x}=(x / g)(d \theta / d t)^{2}+(z / g) d^{2} \theta d t^{2}$- e. Yaw maneuver, lateral left and right test directions. The additional load factor $\left(\Delta \mathrm{N}_{\mathrm{Y}}\right)$ induced by yaw, is computed as follows:$\Delta \mathrm{N}_{\mathrm{Y}}=\left(\mathrm{y} / \mathrm{g}\right)\left(\mathrm{d} \psi / \mathrm{d} \mathrm{t}\right)^{2}+\left(\mathrm{x} / \mathrm{g}\right) \mathrm{d}^{2} \psi / \mathrm{d} \mathrm{t}^{2}$- f. Yaw maneuver, fore and aft test directions. The additional load factor (ΔNx) induced by yaw change, is computed as follows:$\Delta \mathrm{N}_{\mathrm{x}}=\left(\mathrm{x} / \mathrm{g}\right)\left(\mathrm{d} \psi / \mathrm{d} \mathrm{t}\right)^{2}+\left(\mathrm{y} / \mathrm{g}\right) \mathrm{d}^{2} \psi / \mathrm{d} \mathrm{t}^{2}$

where:

x = fore and aft distance of materiel from the aircraft CG, m (in.) y = lateral distance of materiel from the aircraft CG, m (in.) z = vertical distance of materiel from the aircraft CG, m (in.) g = acceleration of gravity, 9.81 m/sec² (386 in/sec²) ϕ = angle of rotation about the X axis (roll), rad d ϕ/d t = maximum roll velocity in rad/sec (if unknown use 5 rad/sec) d ²ϕ/d t² = maximum roll acceleration in rad/sec² (if unknown use 20 rad/sec²) θ = angle of rotation about the Y axis (pitch), rad d θ/d t = maximum pitch velocity in rad/sec (if unknown use 2.5 rad/sec) d ²θ/d t² = maximum pitch acceleration in rad/sec² (if unknown use 5 rad/sec²) ψ = angle of rotation about the Z axis (yaw), rad d ψ/d t = maximum yaw velocity in rad/sec (if unknown use 4 rad/sec) d ²ψ/d t² = maximum yaw acceleration in rad/sec² (if unknown use 3 rad/sec²)