520.5-22 tailoring guidance

2.1.1 Effects of Combined Environments

Temperature, altitude, humidity, input electrical power, and vibration can combine synergistically to produce the following equipment failures. The synergistic effects may induce failures that would not be exhibited during individual environment testing. In addition to unique effects, the synergistic environments may amplify the stress effects when compared to the effects of individual environments, to include power effects. The following examples are not intended to be comprehensive. For additional information refer to the individual methods.

• a. Shattering of glass vials and optical materiel.
• b. Binding or loosening of moving parts.
• c. Separation of constituents.
• d. Performance degradation in electronic components due to parameter shifts.
• e. Electronic optical (fogging) or mechanical failures due to rapid water or frost formation.
• f. Cracking of solid pellets or grains in explosives.
• g. Differential contraction or expansion of dissimilar materials.
• h. Deformation or fracture of components.
• i. Cracking of surface coatings.
• j. Leakage of sealed compartments.
• k. Failure due to inadequate heat dissipation.

1. If the material is powered, component over-temperature failures.

• m. Printed Circuit Card failures due to short circuiting.
• n. Failure of Electromagnetic Interference (EMI) filters.
• o. Errors due to input electrical power frequency variances.

2.1.2 Sequence Among Other Methods

• a. General. Use the anticipated life cycle sequence of events as a general sequence guide (see Part One, paragraph 5.5).
• b. Unique to this Method. Procedure I is intended to be used before final materiel designs are fixed. If vibration is performed separately from the remaining combined environments, vibration shall be performed first.

MIL-STD-810H

METHOD 520.5

2.2.1 Procedure Selection Considerations

The choice of test procedure is determined by the developmental phase of the equipment (Design, Specific Mission, Troubleshooting, or Envelope Verification).

2.2.2 Difference Among Procedures

While all of the procedures cover the same forcing functions, they differ on the basis of test severity, combination of forcing functions, and applicable mission(s) as defined in the following:

Procedures I and III: These procedures encompass the full operational envelope as defined by the equipment LCEP, the platform/equipment specification, and as further defined for the current/projected installation location(s). The stress levels used in Procedure I may exceed the LCEP parameters in order to establish design margins. The stress levels used in Procedure III would apply to a UUT with a more mature design and would incorporate the specification test levels, to include the maximum range of climatic, input electrical power, vibration, and operational modes/conditions.

Procedure II: This procedure is constrained to the climatic, dynamic, and operational parameters defined for a specific geophysical and climatic area/model. This is further defined by the specific micro-environment in which the equipment operates (equipment bay, cockpit, etc.), local electrical power quality, and operational modes. Test results are relevant only for operation within the specified envelope.

2.2.2.1 Procedure I - Engineering Development

Use Procedure I to help find design defects in new or modified equipment while it is still in the development stage. A combined environment test is useful for this purpose since it will reveal synergistic failures. The primary purpose of this test is to uncover any anomaly, with the exact cause being secondary. A root cause analysis is then performed to determine the corrective action. Subsequent testing, to aid in the root cause analysis, may be enhanced by using higher stress levels than the item is likely to encounter on a regular basis in the field. Test duration shall be based on time required to induce expected failure modes, such as those listed in paragraph 2.1.1. Test parameters may be chosen to emphasize specific environmental effects. However, using limited environments and stressing materiel items beyond realistic limits may induce failures that would not occur under realistic conditions. This procedure is not a substitute for Environmental Stress Screening (ESS) or Highly Accelerated Life Test (HALT) testing. Given these cautions, perform the Procedure I steps specified in the Test Parameter Selection/Profile Development Table (Table 520.5-VIII).

2.2.2.2 Procedure II - Flight or Mission Support

This procedure is performed in preparation for a specific mission scenario or functional testing; and also for troubleshooting of fielded materiel exhibiting specific mission problems. This can include issues which only occur during specific combined environment(s) and which may be resolved when these conditions no longer exist. Its purpose is to use laboratory conditions to evaluate and resolve materiel issues pertinent to specific mission scenarios and to resolve issues prior to resumption of flight testing of developmental or fielded materiel. In addition, this procedure can be used where a standard qualification process is not feasible due to deployment schedule constraints; i.e., urgent user needs. These types of programs require performance during specific missions but normally are not required to operate to the full platform envelope.

This test is not accelerated; the damage accumulation in the test should be no faster than in operational or in-flight testing. Therefore, development hardware can be interchanged between laboratory and flight or functional testing. In general, use a test duration representative of the design mission or, if troubleshooting, sufficient to identify materiel

MIL-STD-810H

METHOD 520.5

problem (may be iterative). For troubleshooting it may be necessary to replicate the life cycle history of the materiel. This procedure is not intended to be used in lieu of Procedure III. Perform the Procedure II steps specified in the Test Parameter/Profile Development Table (Table 520.5-VIII).

2.2.2.3 Procedure III - Platform Envelope

The Platform Envelope test is intended to demonstrate compliance with specific platform/equipment specification requirements of combined synergetic environmental conditions. This testing emphasizes the most significant environmental stress conditions and combination of stress conditions; in accordance with the operational envelope as determined by the LCEP, specification, and developmental maturity of the materiel. Include in the platform envelope test the maximum amplitude of each stress and any unique combinations of stress types. Use caution in the application of maximum amplitude combinations that would expose the materiel to unrealistic conditions. The intent is to demonstrate maximum conditions that a specific platform may expose the materiel to. Use a test duration of a minimum of 10 cycles. For power cycling the input power shall be cycled such that the electrical stresses are balanced throughout the test. Cycling shall start at nominal, then high voltage, low voltage, etc.; ending with nominal. Additional cycles may be conducted, but platforms are rarely exposed to their maximum envelope conditions and a failure discovered in subsequent cycles may be caused by unrealistic extended exposures. Perform the Procedure III steps specified in the Test Parameter/Profile Development Table (Table 520.5-VIII).

2.2.3 Selecting Combined Environments

The intent is to identify all synergistic environmental combinations from the LCEP and to develop a representative test profile to verify that the materiel operates properly throughout the LCEP derived environments, within a single test. The relevant environments are temperature, altitude, humidity, input electrical power, and vibration. However, this method can be tailored to perform a series of combined environmental tests providing that all synergistic effects can be evaluated. An example of a series of combined tests would be the performance of a temperature, altitude, humidity, input electrical power test, and vibration. Apply the guidance in section 2.3. Each combined environment test shall be completed in single climatic chamber to maintain materiel conditions and eliminate unintended environmental stresses. Test items shall not be transferred between chambers due to combined environment equipment capability limitations.

2.2.4 Environmental Stresses

• a. Determine environmental stresses including temperature, supplemental cooling, altitude, humidity, and input electrical stresses, and vibration.
• b. Determine test levels (based on measured data, if available, or from the LCEP) for each stress from the mission profile information in the manner described in paragraphs 2.2.4.1 through 2.2.4.5. Other information, such as engine rpm or data on the platform's system environmental control system (ECS) may be needed.

2.2.4.1 Thermal Stress

The thermal stresses that materiel experiences during a mission is dependent upon the localized ambient temperature, induced/contributing thermal influences (solar, reflected radiation, adjacent heat producing equipment, etc.), altitude and humidity conditions in the equipment location. This may be either external or an internal compartment or bay. Contributing factors, such as flight conditions, power requirements, and the performance of supplemental cooling to the materiel are to be considered. (Thermal stresses shall consider steady state conditions, transition temperatures, ramp rates, and extremes for both operational and non-operational conditions.)

• a. Ambient worldwide climatic extremes are found in MIL-HDBK-310. The ground level ambient temperatures are based on data from standard meteorological instrument shelters. They represent free air temperatures in the shade about 1.5 meters above the ground. The materiel ground soak temperatures in each mission are not necessarily directly correlated to the meteorological data. Ground soak temperatures are influenced by the albedo of the material, solar radiation, reflected radiation, heat from adjacent equipment, etc. Temperatures at altitude represent actual extreme high and low measurements, compiled from multiple data measurements. They may not represent conditions at a specific geophysical location. When determining air temperature change rates during ascent or descent conditions, the proper adiabatic lapse rate for the local atmospheric conditions (dry or moist,) shall be utilized. In addition to the highest/lowest recorded values, MIL-HDBK- 310 also includes the 1, 5, 10, and 20 percent frequency of occurrence values. If used in preparing a thermodynamics analysis, then additional forcing functions (direct solar radiation, reflected radiation,

adjacent equipment, etc.) must also be included in the analysis. The materiel thermal response during the transition from ground soak (non-operational/standby) conditions to operational conditions will demonstrate a thermal lag where the equipment temperature will lag the ambient temperature based on the equipment mass, material, and airflow conditions.

• b. The environmental test conditions for any test item are dependent on the local thermal environment; including the type of local cooling (supplemental cooling air, ram air, convective cooling, etc.), induced heating from solar radiation and other thermal radiators (engines, exhaust ducts, adjacent equipment, etc.), and heated air from adjacent convectively cooled equipment. Systems comprised of multiple line replaceable units (LRU) may be tested concurrently in one chamber if the LRUs all have the same local operational environment. Likewise, systems comprised of multiple LRUs shall be tested separately if the LRUs have different local environments.
• c. The thermal stress test parameters to be used in performance of Procedure II - Flight or Mission are derived from the specific mission segments of the selected platform flight profile. For troubleshooting of fielded materiel exhibiting specific mission problems, select the thermal environment of the mission segment (s) in which the problem exists. For mission unique programs, which will operate in specific geophysical areas, select all of the thermal conditions throughout the entire flight profile. Use of measured temperature from the platform is preferable. Use natural temperature conditions derived from standard atmospheric models and induced temperatures due to air friction against the platform exterior and/or solar loading for materiel that is mounted externally to the platform. If data is not available for bay/compartment temperatures, perform an analysis of typical Environmental Control Systems, induced, and natural conditions or obtain/acquire measured data for the area of concern. Many modern electronic avionic equipment record internal component temperatures and can be used to analyze the temperature conditions during testing of the bay or cabin compartment. By adjusting the chamber air temperature conditions with the equipment operating in the same operational mode, match the response temperature of the equipment as it would be in-flight. Choose the climatic region, Basic, Hot, Cold, Severe Cold, and Coastal/Ocean and use the 20% frequency of occurrence temperature for day or night as starting points for ground conditions of start-up, taxi, takeoff, and landing or measured data at the platform sight.
• d. The thermal stress test parameters to be used in performance of Procedure III - Platform Envelope are derived from the equipment/platform performance specifications and the LCEP. For equipment which is operated only during specific segments of the platform envelope, use the thermal parameters corresponding to the equipment’s operational envelope. For equipment that is to be used on multiple platforms, use the worst case parameters for those platforms. Use caution, when applying parameters from multiple platforms, do not apply unrealistic combinations of environmental conditions. The duration of the thermal stresses shall be sufficient for thermal stabilization and representative of the anticipated durations determined by the LCEP.

2.2.4.1.1 Bay Conditions. (Reference MIL-STD-2218)

The effective air temperature within equipment bays must be determined in developing a test profile. Ram air will track the external temperature. Cold air at altitude may induce thermal shocks, and this condition must be considered when developing the test profile. For Environmental Control System (ECS) cooled equipment the ECS specification requirements must be met in regards to the cool down ramps and temperatures. See Table 520.5-I for an F-15 supplemental cooling example.

• a. Ram-air-cooled compartments. Use this section to determine the bay temperature for an avionics or system in a compartment that is ram-cooled. Utilize thermal analysis and/or thermal surveys to determine the effective air temperature in the compartment at specific altitude and temperature conditions. Determine the thermal stress in a ram-air-cooled compartment from the following relationship.

$\mathrm{T}_{\mathrm{eff}}=\mathrm{T}_{\mathrm{amb}}\left[1+0.2 \mathrm{M}^{2}\right]$

where:

T_amb = Ambient air temperature (K) at altitude being flown (Tables 520.5A-I, 520.5A-II, and 520.5A-III).

Teff = Effective air temperature (K) as modified by air velocity cooling effects and used in the test cycle.

M = Mach number being flown

MIL-STD-810H

METHOD 520.5

• b. Environmental Control System conditioned supplemental-air-cooled bay. Use this section to determine the thermal stress for an avionics system located in a bay that receives its cooling from the platform ECS. To the extent possible, the effects of the ECS should be simulated in the chamber. If the ECS cannot be simulated in the chamber, utilize thermal analysis and/or thermal survey data to determine the effective air temperature in the compartment at specific altitude and temperature conditions. Determine the mass flow rate and temperature of supplemental air for each break point in the mission profile. Model the onboard ECS in terms of its primary components such as pressure regulators, heat exchanger, turbo machines, water separator, etc. If the heat load from these systems is significant, include the mass flow rate being injected into the bay and the location of other systems in the calculation.
• c. Materiel supplemental cooling thermal stress. Use this section to determine the effect for test items that require supplemental cooling from the platform. Evaluate the component for additional supplemental cooling not mentioned in the previous section. This cooling may be direct air or liquid cooling into the materiel or through a cold plate. For ECS systems that are open to the external environment, the mass flow rate shall be system temperature and mass flow rate parameters.)

2.2.4.2 Altitude Stress

Use altitude simulation to evaluate the effect of various pressures (high or low) across the platform envelope throughout the LCEP. Method 500 addresses the effects of low pressure as an individual forcing function. Method 520 addresses high and low pressure and its effects in conjunction with high/low temperatures and high/low humidity forcing functions, as well as the effects of the altitude ramp rate. Vibration may also be added as a forcing function.

For airborne materiel the pressure altitude has a direct effect on the thermal performance of the materiel (See paragraph 2.2.4.1.1 for additional information on bay conditions). Convective cooling is directly proportional to the density (mass per unit volume) of the surrounding air. At high altitude the air density is less, thus the air mass available for convective cooling is diminished. This will result in higher operational temperatures. This also applies to equipment receiving cooling air via an ECS. For an open system, where the system is open to the natural environment, the air velocity may be constant; however, the air mass flow rate will be reduced in proportion to the pressure. When performing altitude tests on ECS conditioned equipment care must be taken to match the mass flow rate with the appropriate altitude conditions.

Variations in air pressure will affect both environmental and hermetic seals. Under normal operation the environmental seals will allow the passage of gases into and out of the materiel in order to equalize pressure (breathing). Rapid pressure changes may result in damage to, or complete failure, of environmental seals. Damage to Electromagnetic Interference (EMI) seals may also occur. Hermetic seals may rupture due to pressure variations. The air mass internal to the hermetic seal remains constant; however, as the external pressure drops, the pressure differential may result in the seal being ruptured. Altitude ramp rates must be controlled to avoid an over test or under test. Unrealistic ramp

rates will result in unintended results. The ramp rates shall comply with the maximum platform altitude ramp rates. Ramp rates that do not meet these requirements will result in an invalid test.

Altitude stress in conjunction with humidity is highly dependent upon the atmospheric model(s) (reference Annex A, Tables 520.5A-I thru 520A-III) in which the materiel is to be operated during its life cycle. For world wide deployment the combined worst case parameters shall be used. See paragraph 2.2.4.3 for the description of the synergistic altitude /humidity effects.

Changes in atmospheric pressure and air density may also effect the electrical characteristics of components and between components. Changes in air density will result in a variable air dielectric field, resulting in changes in capacitance between and within components. For high voltage materiel this may result in corona effects and arcing. Some types of capacitors may swell due to decreased air pressure, resulting in changes to their capacitance.

The test site and test chamber capabilities must also be considered when performing altitude tests. Most materiel specifications require that the materiel operate from Mean Sea Level (MSL) to some maximum altitude. However, most altitude test chambers are designed such that the maximum pressure corresponds to the site ambient pressure. For pressure sensitive material (altimeters, air data computers, pressure sensors, pressure switches, etc.), where it is required that testing be performed across the full operational pressure envelope, a test site, or test facility, shall be chosen that is capable of meeting the full pressure envelope of the material.

Reference U.S. Standard Atmosphere (1976)

MIL-STD-810H METHOD 520.5

2.2.4.3 Humidity Stress

Absolute humidity is the mass of water vapor in a specified volume of air. It may be expressed in many ways, but is generally specified as grams/m3 or parts of water vapor per million part of dry air (ppm). Data for the highest and lowest worldwide absolute humidity is found in MIL-HDBK 310, along with data for absolute humidity at altitude. The dew point, the temperature at which condensation would occur if the air was cooled at constant pressure, is the observed meteorological element used to calculate the absolute humidity. Relative humidity is the ratio of the current absolute humidity to the highest possible absolute humidity (which depends on the current air temperature). A reading of 100 percent relative humidity means that the air is totally saturated with water vapor and cannot hold any more. For a given volume of air, with a given absolute humidity, the relative humidity will increase as the temperature decreases until the dew point is reached. At that point the water vapor will condense out as liquid water.

The effects of ambient humidity on materiel depends on the humidity levels, temperature, pressure, and duration of exposure. For long term exposure effects see MIL-STD-810, method 507. The effects of interest in this test method pertain to effects due to condensation and freezing of water vapor when the dew point is reached during materiel operation and also effects due to absorption and intrusion of water within material. Note that the materiel temperature may be lower than the ambient environment due to black body radiation.

When deployed in a Warm/Moist environment, materiel and the materiel enclosures (equipment bays, cockpit, cabins, etc.) will experience the absolute humidity of the ambient environment. Upon transition to a cold environment (aircraft climb to altitude) the water vapor will condense when the temperature drops to the dew point. This may result in liquid water pooling within electronic equipment; with the potential to create short circuits, sneak circuits, and other adverse effects. The water may subsequently freeze. The expansion of water during freezing may damage components, degrade seals, and lead to delamination of material.

Humidity stresses also occur during transition from a Cold/Dry environment to a Warm/Moist environment (aircraft descent from altitude). In this case the materiel temperature transition lags the transition of the ambient environment. Humidity from the Warm/Moist environment will then condense on, and within, the cold surfaces of the materiel. This may result in liquid water pooling within electronic equipment; with the potential to create short circuits, sneak circuits, and other adverse effects. If the materiel surface temperature is below the freeze point the water may subsequently freeze. The expansion of water during freezing may damage components, degrade seals, and lead to delamination of material.

During environmental testing the intent is to simulate the ascent and descent of the aircraft within, and through, the varying atmospheric temperature and humidity levels. During transitions from high altitude /cold to low altitude/warm/moist, the humidity ramp rate shall be sufficient to induce moisture condensation and frost on the test material.

For materiel located within an ECS controlled environment, the efficiency of the water separator will be an important consideration during the test. ECS operation may result in high relative humidity (RH) due to chilling of the air. This will produce condensation if the dew point is exceeded. Characterization and replication of the ECS system is essential in the development and validity of testing materiel in this environment. When the efficiency of the ECS is unknown, use the approximation technique in 2.2.4.1.1 b.

For this test, whenever the cold day environment is being simulated, humidity will be uncontrolled, but less than or equal to the dew point temperature in Annex Table 520.5A-II For the hot environment, dew point temperatures will be less or equal to values in Annex Table 520.5A-I. For the Warm/Moist day, dew point temperatures will be greater than or equal to the values in Annex Table 520.5A-III up to 10 km (6.2 mi) altitude. Above 10 km (6.2 mi), the dew point temperature is less than or equal to the values in Annex Table 520.5-II. If the platform has an ECS, the design specifications for the Warm/Moist day apply.

2.2.4.4 Electrical Stress

Electrical stresses are expected deviations of the materiel's electric supply parameters from their nominal values at the materiel terminals. Every aircraft electrical power system is required to be designed to ensure the retention of the electrical characteristics as specified in the selected standard document, throughout the full range of operational and environmental conditions likely to be encountered in the aircraft in which it is installed. (Reference STANAG 3456 Edition 7). The test procedure must simulate to the required extent, all electrical stresses occurring during normal operation in service (mission profile) that contribute synergistically to the environments. In addition, appropriately demonstrate operation of the test materiel's functions at each test condition. It is not the purpose of this test to simulate

MIL-STD-810H

METHOD 520.5

extremes specified for special situations or to take the place of electrical stress tests. Simulate special conditions such as emergency operation of certain aircraft materiel within the electrical/electronic system only on request. Depending upon the requirements and the availability of data, the simulation may cover the range from the exact reproduction of the specific electric supply conditions within a special aircraft for a specific mission profile, down to a standardized simplified profile for generalized applications.

A significant percentage of the materiel life cycle occurs during maintenance operations, which may use ground power from ground support equipment. Ground support equipment normally replicates platform power, but input power measurements may be necessary if the equipment is sensitive to input power anomalies.

Consider the following conditions and effects to determine whether they affect the operation and reliability of the materiel to be tested. See Part One, Section 5.1.c, Test Conditions, Input Power for guidance to available sources of voltage and frequency characteristics or Tables 520.5-IV, 520.5-V and 520.5-VI that contain steady state voltage and frequency conditions.

• a. AC system normal operation stresses.
• b. Normal ON/OFF cycling of materiel operation.
• c. DC system normal operation stresses.
• d. Electrical stresses induced by mission-related transients within the electrical system.
• e. Abnormal steady state stresses of voltage and frequency if the material is required to operate for safety of the mission or personnel.
• f. Emergency steady state stresses of voltage and frequency if the material is required to operate for safety of platform or personnel (flight critical).

2.2.4.4.1 AC & DC System Normal Operation Stresses

Voltage variations are quasi-steady changes in voltage from test cycle to test cycle. A suggested input voltage schedule would be to apply the input voltage at platform nominal voltage for the first test cycle, at the platform high normal voltage for the second test cycle, and at platform low normal voltage for the third test cycle. This cycling procedure would be repeated continuously throughout the test, with the last cycle performed at the nominal voltage. However, if troubleshooting, then use the power characteristics at which the failure or anomaly occurred, along with the other environmental conditions at the time of the anomaly.

For AC systems, the input electrical power frequency variations shall also be considered when developing the test cycle. The input power frequency will vary during normal operation due to variations in the generator RPM and platform power loads. The frequency variations may result in increased thermal loads leading to equipment failures or abnormal operation of equipment. Some failures, such as the failure of EMI filters, may not be obvious. This should be considered in developing the overall test development schedule.

See Tables 520.5-IV, 520.5-V and 520.5-VI for normal input power parameters. Additional data may be found in MIL-STD-704, RTCA-DO160, and MIL-STD-1275. For shipboard power refer to MIL-STD-1399, Part 300.

A suggested input voltage and frequency variation schedule is provided in Table 520.5-III.

Note 1. Refer to MIL-STD-704 for additional information concerning electrical power characteristics.

• Note 2. The MIL-STD-704 version may have additional notices. Refer to the latest notice to find any changes. Most notices do not vary the input power steady state conditions.
• Note 3. ***360 – 457 Hz (V-22) Special Frequency Range, see V-22 platform specification.
• Note 4. If material is to be used on multiple platforms, use the lowest and highest levels of all applicable MIL-STD-704 versions.

Note 5. Refer to STANAG 3456, ISO 1540, or EN 2282 for NATO member aircraft electrical power characteristics.

• Note 6. For the C-130, F-22, and U-2 aircraft platforms, contact the appropriate Platform Program Office for electrical power conditions.

MIL-STD-810H METHOD

2.2.4.4.2 Normal ON/OFF Cycling of Materiel Operation

Turn the materiel on and off in accordance with materiel operating procedures outlined in appropriate technical manuals, to simulate normal use.

2.2.4.5 Vibration Stress

Method 514 contains guidance for characterizing the vibration environment through measured data and laboratory vibration specification development, with Operational Tailoring Guidance provided in Annex D. Default vibration profiles are provide in Method 514, Annex D, for use when measured data are unavailable. Caution should be used when applying the default vibration profiles in Method 514 for functional tests since the default profiles often contain significant time compression and conservatism. As a result, the default vibration levels may not be representative of the functional vibration environment. When the time compression factors are known, adjustments to the default profiles may be made by removing the time compression factor using the Miner-Palmgren Equations (i.e., Miner’s Rule) discussed in Method 514, Annex F, paragraph 9.2.1. The resulting levels will typically be more representative of the maximum service levels for use in the functional vibration test. Reference Method 514, Annex A, paragraph 2.1.2.1.

Note that conservatism factors in the Method 514 default vibration profiles are imposed for various reasons (see Method 514, paragraph 2.1b) that are applicable to the functional vibration test. Thus conservatism factors should not be removed from the default vibration profiles without careful consideration.

For Procedure I the vibration levels shall replicate the materiel design/specification levels. For Procedure II the vibration levels shall be based on a particular mission profile. For Procedure III the vibration levels shall be based on the maximum performance envelope of the aircraft.

• a. The vibration stresses to be considered for the test cycle are those due to both attached and separated aerodynamic airflow along the vehicle's external surfaces, jet engine noise, or pressure pulses from propeller or helicopter blades on the aircraft structure. Determine the vibration spectrum and level for each mission segment by careful use of measured data. Apply the guidance written below in those cases.
• b. In many instances, field/fleet flight data are not available for the specific aircraft, materiel location in the aircraft, or flight phases. In such cases, there are several analytical techniques for vibration, spectrum, and level prediction that can be used to determine vibration test conditions (see Method 514 and Table 520.5- VII).
• (1) Scaling vibration test conditions from data obtained on another platform at a different materiel location, or for a different flight condition has to be done with extreme care because of the numerous nonlinear relationships involved and the limited amount of data being used. For example, maneuver-induced vibration conditions generally cannot be predicted from cruise vibration data. A more prudent approach is to use the linear dynamic pressure models in Method 514.
• (2) In all cases, field/fleet flight vibration data should be defined in accordance with Method 514 following the vibration characterization guidance in paragraph 2.3.
• c. Because of the nature of vibration control equipment, it may be difficult to change vibration level and spectrum shape in a continuous, smooth manner. Therefore, the mission profile may be divided into segments over which it will be assumed that the vibration level and spectrum shape is constant for test purposes. In addition, vibration specifications are typically defined in three orthogonal axes and the vibration tests are typically conducted in three sequential axis vibration tests, unless a multi-axis test apparatus is available. Ideally, each segment of the mission profile should be divided by the three axes of vibration. However, in a combined environment, changing test axes in the middle of a segment of a mission profile may be impractical and may invalidate the test results. In this case, it may be necessary to perform the functional vibration only in the worst case axis in terms of vibration level and materiel sensitivity. Alternatively, different axes may be tested in different phases of the test cycle.
• d. Unless field/fleet data exist, the appropriate tables and figures of Method 514, Annex D, are used to determine vibration conditions except as modified in Table 520.5-VII.

MIL-STD-810H METHOD

2.3.1 Test Cycle Development

A test cycle is defined in this method as a series of test segments simulating different climatic/dynamic/power input conditions. In general, a test cycle is composed of separate temperate/altitude/humidity segments: cold/dry, cold/dry/altitude, warm/moist, hot/dry, hot/dry/altitude, etc. Additional test parameters (power, vibration, supplemental cooling air, etc.) are incorporated as required to simulate the operational environment. For engineering development (Proc. I) and platform envelope (Proc. III) testing, the profile is defined as a set of life cycle conditions encompassing the materiel design/specification requirements and arranged to encompass all expected combinations of potential mission environments (Figures 520.5-3a, 520.5-3b and Table 520.5-VIII). For mission specific testing, a profile is defined as a performance-environmental condition-time history of the specific mission of concern. A specific mission profile (Proc. II) may be divided into segments such as takeoff, cruise, combat, maneuvering, landing, maintenance, etc., with each segment replicating the climatic/dynamic/power input conditions representative of that mission segment (Figure 520.5-2).

When performing Procedures I or III, the test parameters are derived from the platform/equipment performance specifications, and design documentation. Using the maximum design parameters, compile a table of the climatic and induced environments, and the combination of those environments that the equipment will be subjected to during its life cycle. Develop a test cycle that simulates each of the appropriate combined environments derived from the LCEP. Each segment of the profile should be of sufficient duration for the environmental stresses to stabilize, and to allow for performance verification.

When performing Procedure II, the test parameters are derived from the specific mission parameters as per the portion of the LCEP that defines this specific mission. Use of measured platform data, if available, from the specific equipment location is essential. If measured data is not available, parameter data may be obtained from the program management office, user needs statement, etc. Identify the specific climate(s) under which the equipment is intended to operate; hot/humid (tropical), hot/dry (desert), cold/dry (desert), cold (arctic), etc. Determine the specific environmental conditions, and combinations of conditions, to which the material will be exposed during performance of the mission. This includes both climatic environments (temperature, altitude, humidity) and induced environments (vibration, power, supplemental cooling). Using this data develop a mission profile inclusive of taxiing, take-off, cruise, combat, maneuvering, landing, maintenance, etc.

When Procedure II is to be used for troubleshooting, then the test parameters are derived from the specific mission conditions under which the equipment experienced a failure. This shall include performance data and measured climatic/dynamic/power data obtained during the timeframe of the failure, to the extent that the relevant data is available. Anomalies experienced during flight operations may be specific to the environmental and dynamic conditions at the time of the anomaly. The anomaly may not manifest itself upon return to ambient conditions. Due to the cumulative and synergistic effects of the various environments, the full mission profile is recommended for initial troubleshooting. As troubleshooting progresses, specific segments of the profile may be exercised to isolate and define the root cause of failure.

Refer to Table 520.5-VIII for guidance on test parameter selection. Reference the LCEP.

MIL-STD-810H METHOD