ии. и-о i и-о lyjt, 14 JULY 18S9

considered for a test item that is varied. Altitude simulation may be considered for a test item that is hermetically sealed, uses pressurized cooling paths to transfer heat, has components that contain a vacuum, has voltages of sufficient potential to arc in the presence of rarefied air, long range missions, or for other appropriate cases. Cooling airflow is required for all test items that use sv4Pplementary airflow in the aircraft.

1-3.2.2.3 Mission profile selection. The first step in constructing a combined environment test is to select the mission profiles to be used. An individual aircraft is designed to operate within a specified flight envelope (Mach number/altitude regime) and to fly specific mission profiles. (Senerally, an aircraft can fly many different missions, such as training, air superiority, interdiction, ground support, etc. In addition, aircraft are flovsi under specialized conditicr.s that simulate a high-threat combat environment. These wartime skill exercises, such as Red Flag, are designed to train operational squadrons under realistic wartime conditions.

usual iy. not ail the missions flown by the aircraft need to be included in the test cycle. It is possible to identify two or three of the most highly utilized mission profiles that, as a group, reasonably approxinate the aLggregate effeefc of all the missions flow by the aircraft. This will adequately simulate the routine deployment life. In addition, the utilization of wartime skill exercises as part of the mission profile will stress the equipment under simulated conbat conditions. To select the mission profiles to be liSed, the following approach is recomnended.

a= Identify all aircraft missions agid the utilisation rate of each mission of the aircraft in which the equipment is to be instal led. This information may be obtained from the operational conmands or the flight manual uaad by aircraft crews. For aircraft under development, the design flight envelopes, design mission profiles, and the design utilization rate of each mission shall be used wien aci/Uax 1 J. igiit uav/a ape nou avaxiauie.

b. Determine the missions that conprise a majority (if possible, 80 percent of total flown) of the total routine, daily mission utilization. To do this, examine the utilization rates for all mission profiles of the aircraft and rank them in order from highest to lowest. Then, take the mission profiles that conprise the majority utilization rate and use these as mission profiles for combined environment testing. Missions with similar functions and flight characteristics can be lunped together to minimize the пглпЬег of profiles to be generated. Table 520.1-1 shows an exanple distribution of missions.

o. In order to simulate the high-threat environment, missions flown under the wartime skill exercises shall be separately identified. These data may be obtained

from the op€ratior.al cofiinsuid or prcvided by the procuring agency.

Once these data have been obtained, two separate test cycles can be constructed according to 1-3.2.2.2. One test cycle using the mission profiles in 1-3.2.2.3b will be developed to simulate routine usage and another test cycle using the mission profiles in 1-3.2.2.2c will be developed to simulate usage under combat or combat-training conditions. METHOD 520.1

TABLE 520.1-1. Exanple utilization, rates, of mission profiles. ! Mission ; Percent Utilization Rate

Groimd Attack, Training : 40

Ground Attack, ConJbat 20

Defensive Maneuvers 20

Seai>ch and Rescue ! 10

Fimetional Check : 5

Training Cycle ! 5

100

Obtain the altitude and Mach nunjber versus time values for each mission profile selected, as ahovn schematically in figure 520.1-3. These parameters of the mission profile are used to calculate the environmental stresses.

1-3.2.2.4 Environmental stresses. The second step is to determine environmental stresses including vibration, tenperature, supplemental cooling, humidity, altitude, and electrical stresses. Test levels for each stress are determined from mission profile information in the manner described in 1-3.2.2.5 through 1-3.2.2.9. Other infors tion, such as engine rpm or data on ths aircrafts environmental control system (EGS) may also be needed.

Since the first three missions, as a group, total 80 percent of the utilization rate, then these three mission profiles would be selected for combined environment testing. If any of the other missions are determined to include extreme or sustained environmental conditions not encountered in the first three missions, then those missions containing these extreme or sustained conditiQn.s and adding the most diversity to the test cycle also should be selected. If the first mission selected is utilized twice as nuch as the other two missions, then Mission 1 should be run twice as much per cycle.

1-3.2.2.5 Vibration stresa. Random vibration shall be applied to all equipment items designated for jet aircraft installation. Random vibration or sine sxjperimposed on random vibration should be used for all equipments designated for propeller aircraft. Vibration of an appropriate level and spectrxJtn shape shall be applied continuously dtiring mission profile simulation in the test cycle. Unless measured data exist, it is recommended that the appropriate tables and figures of method 514 be used to determine vibration conditions except as modified in table 520.l-II.

Short duration vibration events and those that occur infreqrjently need not be included in the test cycle. These events include firing of onboard guns, general aircraft motion, and shock of hard landings. These events па,у be tested separately using the appropriate test method.

1.00-1

n TC

0.?5 Ч

0.00

30.00 H

15.00 -I

7.50 H

0.00

0.00

Altitude (kft)

-I-1---1-

6C.O0 120.00 180.00

Time (min)

300.00 360.00

FIGURE 520.2-3. Schematic aiasion profile, altitude and maeh number.

ftETBOD 520.1

TABLE 520.l-II. Suggested random vibration test eriteria for aircraft equipment,

JET AIHCRAFT

Use table 514.3-III with these modifications:

К = 6.8 X 10~® for cockpit panel equipment and equipment attached to structure in conpartments adjacent to external stjrfaces that are smooth, free from d i scon11nult i es.

к = 3.5 X 10 for equipment attached to structure in conpartments adjacent to or Inmediately aft of external surfaces having discontinuities (cavities chines, blade antennas, speed brakes, etc.) and equipment in wings, pylons, stabilizers, and fuselage aft of trailing edge wing root.

If Mach nuntoer is not in the range of 0.85 to 0.95 the calculated levels can be reduced by 5 dB.

For propeller aircraft and helicopters, use appropriate tables in method 514.4

For those segments with the same vibration spectrum shape, the following analysis can be used to reduce the nunber of vibration test levels. The discussion is in terms of the suggested spectrum shapes for jet, rotary wing or propeller aircraft of method 514.

For test purposes a vibration level for each mission segment can be determined using the altitude and Mach nuiber plots for each mission. (Note: For test purposes the larger of due to aerodynamic or due to jet engine noise is utilized at any point in time in the mission.) The maximum value that occurs in each mission shall be identified. All segments of the mission that have values within three dB of the maxinun shall be considered, for test ptirposes, as having a constant value determined using the value of *дх~* segments of the mission that have

dyr mic pressure values between ofUf-Z dB and Wjijx - - all be considered for test purposes as щдх - having a constant value determined using the value of oMAX dB. This process of identifying three-dB bands of dynamic pressure values, over wtiich is considered to be a constant and whose value is determined by using the dynamic pressure values of the batnds midx>int. is continued until the calculated value is less than 0.001. For test purposes, segments of the mission with calculated values of less than 0.001 can be set equal to 0.001 unless the test facility can control below this test level.

The value of VU, reflects the changes in aerodynamic flow field around the aircraft. A cruise Wj value reflects normal angle of attack flight, while a maneuver value reflects highly separated flow conditions which induce intense low-frequency aircraft vibration.

TABLE 520.1-III. Ambient outside air tenperatures.

№ГНОО 520.1

520.1-10

Mlb-STD-SiGE 14 JULY 1388

TABLE 520. l-III. Airfcient outside air tenperatures (eontinxjed) . МШШ NDIST ATNDSPHEBE MDDEL:

1/ Uncontrolled humidity (dry as possible)

2/ Qround soak tenperatures are not necessarily related to measured data but are extreme levels to reduce ground soak time.

The vibration stresses to be considered for the test cycle are those diie to both attached and separated aerodynamic airflow along the vehicles external surfaces. Jet engine noise, or pressure pulses from propel ler or helicopter blades on the aircraft structure. Ths vibration spectrum and level can be determined for each mission segment by careful use of measured data. Guidance sritten below shall be applied in those casea.

In many instances, measuired flight data are not available for the specific aircraft, equipment location in the aircraft, or flight phases. In much cases, there are several analytical techniques for vibration spectrum and level prediction that can be used to determine vibration test conditions (ref. a).

The scaling of vibration test conditions from data Bieasmed on wiother aircraft, at a different equipment location, or for a different flight condition has to be done with extreme care because of the numerous nonlinear relationships involved and the limited amsTjnt of data being utilized. For exanple, maneuver-induced vibration conditions generally cannot be predicted from cruise vibration data. A more prudent approach is tc utilize the linear dyr.amic pressure medals in usthod 514.

In all cases, measured flight vibration should be in acceleration power spectral density (PSD) format based on one-third octave analysis or 20 Hz or narrotver constant-bandwidth analysis. Experience has shown that the use of a standardized vibration spectr-tsn shape and the modified levels of method 514 yield as good results in terms of equipment deficiencies as the use of the highly shaped vibration spectra (ref. b).

Becaiise of the nature of vibration control equipment, it is difficult to change vibration level and spectrum shape in a continuoue, smooth manner. Therefore, the mission profile has to be divided into segments over which it will be assuned that

the vibration level агк1 spectrum shape is ccristant for test purposes.

1-3.2.2.6 Bav-thermal stress. The thermal stresses that internal ly-carrled avionics equifMuent experiences during a mission are dependent upon the ambient conditions, fllt conditions, and the performance of the ECS. For the purposes of this test, the ambient outside air conditions shall be as shown in table 520.1-111 for the hot, warm moist, and cold day environments. Hot and cold arabient envlronssnts of table 520.1-III are based on the 20 percent worldwide climatic extreme envelopes from MIL-STD-210. The warm moist environment is based on the tropical environment shorn in MIL-STI>-210. These temperature values are to be used as the anbient conditions for themodynanic analyses for the development of the mission profile test conditions. The ground soak tenperatures in each mission are not necessarily related to measured data. The values shown in table 520.1-III are extreme conditions that have been used in previous programs to accelerate time and reduce tims bstvssen transitions from ons mission to another.

The specific environmental test conditions for anytest item are dependent on the type of cooling for the conpartment in vvhich the eqiiipment is to be located (air-conditioned or ram-air cooled). Avionics equipment systems that consist of more than one black box nay require different snvironmsntal test CKMidltlons for each black box. (For exanple, vtien boxes are in different aircraft conpartments.) For the ссязяэп case of two-black-box system where one box is cooled by supplemental air or fluid and the other box is amblently cooled, both boxes can be tested in one chanber as long as appropriate vibration and altitude slaulation for each box can be achieved. The thermal stimulation would be realistic since the anblent-cooled box would respond to the anbient teaperature similation wile the box that required sivplementai cooling would be primarily responsive to the supplemental cooling air or fluid.

For the ptrposes of this test, th€ following type of thsrmodyr.amic su .aly3is is adequate. A гяоге detailed analysis can be utilized, if desired.

The mission profile time history of altitude and Mach nunnber from 1-3.2.2.3 is analyzed to identify each break point at which the slope of either the altitude or Mach тжпЬег plots change. A thermodynamic analysis is done at each break point using steady-state thermodynamic relationships. Between each break point, linear interpolation is done on each stress to eonstr-uet a continuous profile for each environmental stress. At each such break point, the thermal stress conditions for a test shall be determined in accordance wnth 1-3.2.2.6.1 and 1-3.2.2.6.2.

1-3.2.2.6.1 Ram-cooled compartments. This section is to be used to determine the bay temperature for.an avionics system in a conpartment that is ram-cooled. The thermal stress in a raun-air-cooied compartment can be determined from the fol lowdng relationship.

T = T [1 * 0.18 M]

where T antoient air tenperature at altitude being flown in degrees Kelvin from table 520.l-III M = Mach nunber being flown

1-3.2.2.6.2 Sxjppleaent-al-air-cooled bay. This section is to determine the bay tenperature for an avionics system located in a bay that receives its cooling from the aircrafts ECS. The mass flow rate and tenperature level of sujjplemental air needs to be determined at each break point in the mission profile. The onboard ECS is modeled in terns of its primary conponents such as pressure regulators, heat excsiangers, turbomacOiinery, water separator, etc. Also, calculate the nass flow rate being injected into the Iwf and the location of other systems in order to determine if the heat load from these systems should be considered (refs. с and d). The calculation of the bay tenperature stress can be done using the following sinplified thermodynamic analysis.

a. Assinne that steady-state thermodynamic relationships are valid.

b. Asstss constant but nominal or typical efficiency constants that can be achieved from good design practices for ьш-ЬотасЫпегу amd heat exehangera,

c. Neglect secondary effects in conponents of ECS (i.e., pressure losses in heat exchanger, tenperature losses in ducts).

1-3.2.2.6.3 EoruipmBnt supplemental thermal stress. This section is used to determine the thermal and mass flow for an avionics system that requires forced or supplemental cooling from the aircraft. 1-3.2.2.6.2 recommends an approach to determine the bay thermal stress for an avionics system located in a supplementally cooled conpartment. This same approach is recommended here writh one aiddition: continue the thermodynamic analysis to determine the tenperatxjre and mass flow

being injected directly into the avionics system. The same sources leed to obtain the information for 1-3.2.2.6.2 are also applicable here.

1-3.2.2.7 Hvjniditv.ptresS. The hjsnidlty stress that an internally carried avionics system experiences is dependent upon the airbient hunidity conditions and the performance of the water separator of the environmental control. (5ome aircraft do not cool equifHiisnt with ECS air, thus the equipment sees only anbient hunidity conditions.) For the purposes of this test, whenever the cold day environment is being sinulated. humidity wri 11 be uncontrolled, but less than or equal to the dew tenperature shown in table 520.1-III. For the hot environment, dew tenperatures will be less than or equal to values shown in table 520.1-III. In the case of the warm moist day, dew tenperatues will be greater than or equal to the values found in table 520. l-III up to lOkm Above 10km, the dew tenperature shall be less than or equal to the values found in table 520.1-111. If the platform has an EGS, the design specifications for the water separator Shall be used to define humidity conditions for the warm moist day. When the efficiency of the ECS is unknonwi, the approximation technique put forth above should be used.

ОТЕ: The forietion ef free water on the test items during c-onbined snvircrimsnt testing car. be a ncrrrsil condition.. It wall occur wiieneyer the te!!perat<JB> of the test item is cooler than the dewpoint tenperature of the air being delivered by the ECS or from ram airflow. This is normal and a realistic condition.

1-3.2.2.8 Altitude stress. Altitude simulation should be enployed *ien there is reason to believe that system performance may be affected by variations in air pressure. Exanples of such situations are: hermetically sealed units that use presstjpized cooling parts to maintain sufficient heat transfer, vacuum conponents where the seal is maintained by air pressijre, and units lAiere change in air pressxjre may caxjse arcing or change of conponent values. When altitude effect is to be tested, the altitude stress, or reduced atmospheric pressijre variations, shall be applied according to the mission profiles selected for test. The altitude, or reduced pressure, is initially applied at the simulated aircraft takeoff and continues at the ргеяеш е changes corresponding to the various flight phases from climb-out to landing. The rate of change of pressxire should reflect the clinb or descent rate of the aircraft while performing the various flight mission phases. The maxintsn pressure (mininum altitude) used for the test shall be that of grouxl elevation at the test site.

1-3.2.2.9 Electrical stress. Electrical stresses are deviations of the equipments electric supply parameters at the equipment terminals from their nominal values. The test procedure nust assure that all electrical stresses occurring during normal operation in service (mission profile) are simulated to the required extent.

It is not the purpose of this test method to siraulate extreaBs specified for special situations or to take the place of special electrical Stress tests. Special conditions, like emergency operation of certain aircraft equipment within the electrical/electronic system, shall be sinulated only on request.

Depending tjpon the requipements and the availability of data, the sinulation may cover the range from the exact reproduction of the specific electric supply conditions within a speeial aireraft for a specific laissicn profile, dcvn tc a standardized sinplified profile for generalized applications. The following conditions as*u effects mmt be taken into consideration to determine whether they affect the operation and reliability of the equifxnent to be tested.

a. AC system normal operation stresses.

b. nofml ОМ/(ЖЕ eling af equipment operation.

c. DC system normal operation stresses.

d. Electrical stresses induced by mission-related transients within the electrical system.

1=3.2.2.9.1 AC gygtea nornal operation stresseg. Voltage variations are quasi-steady changes in voltage from test cycle to test cycle. Input voltages shall be maintained at 110 percent of nominal for the first test cycle, at the nominal for the second test cycle, and at 90 percent for the third test cycle. This cycling procedure is to be repeated continuouaily throughout the test. However, if a failure is euspected, this sequence may be interriqjted for repetition cf input voltage conditions.

) 1-3.2.2.9.2 normal Ojl/OFF cvcl in< of equipment operation. The equipment shal 1 be turned on and off, in accordance with equi{xiient operating procedures outlined in 4>propriate technical manuals, to sinulate normal use.

1-3.2.2.9.5 DC gy-stem noFiaal operation str-esses.

a. Voltage variation. See 1-3.2.2.9.1

b. Ripple volta<e. Ripple is the cyclic variation about the mean level of the DC voltage dialing steady-state DC electric system operation. Values shal 1 be taken from actual flight data or from the applicable system specification if flight data are not available. Ripple voltage shall be applied continuously dxiring the mission eimulation portion of each teat cycle.

1-3.2.2.9.4 Electrical ptreaaea ,induced, py niaaion-related ..tranaienta..within the electrical avatera. Unleaa the equipment haa its own power aiply which ia not affected by the transients mentioned, or the equipment is not influenced by

these electrical stresses at all, these stresses aust be reproduced during test. Thm reproduction has to cover all tramslents - like power eur>ges, voltage peaks, electrical current changes, phase unbalance, etc. - which may influence the equipment on teat and are induced by the mission-related operation of the aircrafts electrical/electronic equipment taken as a whole (switching equipment on or off, operating with changing power output, short-time system overload, diiiering generator rpm. operation of regulating devices, etc.).