CN104007351A - Method for determining spacecraft electronic assembly heat cycle test scheme - Google Patents

Abstract

A method for determining a spacecraft electronic assembly heat cycle test scheme comprises the first step of determining a heat cycle test alternative scheme, the second step of calculating the temperature difference delta T1 of all components of a spacecraft electronic assembly under the heat cycle test condition, the third step of analyzing the heat cycle test validity of the spacecraft electronic assembly, the fourth step of calculating the temperature difference delta T2 of all the components of the spacecraft electronic assembly under the normal work condition, the fifth step of analyzing heat cycle test damage of the spacecraft electronic assembly, the sixth step of analyzing normal work damage of the spacecraft electronic assembly, the seventh step of analyzing heat cycle test acceptability of the spacecraft electronic assembly and the eighth step of finally determining the heat cycle test scheme. According to the method, the heat cycle test scheme can be prevented from causing over-tests and under-tests on the spacecraft electronic assembly, and the application risk of the spacecraft electronic assembly in the astronavigation project is lowered.

Description

A kind of Spacecraft Electronic component heat cyclic test scheme is determined method

(1) technical field:

The present invention relates to a kind of Spacecraft Electronic component heat cyclic test scheme and determine method, whether be applicable to evaluate different Spacecraft Electronic assemblies adopts the scheme of various thermal cycling test schemes rationally effective, be devoted to avoid thermal cycling test scheme to cause overtesting and undertesting to Spacecraft Electronic assembly, reduce the application risk of Spacecraft Electronic assembly in space technology, belong to spacecraft thermal cycling test technical field.

(2) background technology:

In Spacecraft Electronic product component development process, reasonable effectively thermal cycling test, can screen out the potential initial failure of Spacecraft Electronic assembly, guarantees reliability and the serviceable life of Spacecraft Electronic assembly.

In engineering reality, Spacecraft Electronic component heat cyclic test Main Basis GJB 1027-2005 < < vehicle, Upper Stage, spacecraft testing require > > and carry out at present.But in GJB 1027-2005, for different Spacecraft Electronic product components, adopt identical testing requirements, the defect characteristics of concrete Spacecraft Electronic product component and working section are lacked to effective specific aim.In Practical Project, also constantly there is the case that the Spacecraft Electronic assembly by thermal cycling test breaks down within cycle designed life.

On the other hand, conventionally analyze thermal cycling test scheme, only consider thermal cycling test scheme validity, the number of faults exposing in test and the ratio of total failare number, do not consider the unnecessary damage that product causes due to test overstress.Analysis deterrmination thermal cycling test scheme, must consider thermal cycling test to the undertesting of Spacecraft Electronic assembly and overtesting risk, to obtain reasonable and effective thermal cycling test scheme.

(3) summary of the invention:

1, object: the present invention seeks to: provide a kind of Spacecraft Electronic component heat cyclic test scheme to determine method, for concrete Spacecraft Electronic assembly, consider undertesting and overtesting, to determine rationally effectively thermal cycling test scheme.

2, technical scheme: the present invention be take thermal cycling test alternatives as research object, for concrete Spacecraft Electronic assembly, by the poor calculating of each device temperature of electronic package and the shear strain of solder joint plasticity, calculate, utilize Coffin-Manson thermal fatigue life theory and Miner linear cumulative damage law, based on thermal cycling test efficiency analysis and the acceptable analysis of thermal cycling test, set up Spacecraft Electronic component heat cyclic test scheme and determine method, it comprises the steps:

Step 1: the determining of thermal cycling test alternatives: according to the requirement of Spacecraft Electronic component design, tentatively determine thermal cycling test alternatives;

Step 2: each components and parts temperature difference Δ of Spacecraft Electronic assembly T under thermal cycling test condition ₁calculate: collect Spacecraft Electronic assembly relevant information, utilize business software, set up Spacecraft Electronic assembly digital prototype, according to the definite thermal cycling test alternatives of step 1, calculate the poor Δ T of each device temperature of Spacecraft Electronic assembly under thermal cycling test condition ₁;

Step 3: Spacecraft Electronic component heat cyclic test efficiency analysis: determine Spacecraft Electronic assembly initial failure device solder joint useful area S ₁, according to the poor Δ T of each device temperature of Spacecraft Electronic assembly under step 2 thermal cycling test condition ₁, calculate Spacecraft Electronic assembly initial failure device solder joint shear strain Δ γ ₁, utilize Coffin-Manson heat fatigue model, calculate Spacecraft Electronic assembly initial failure device thermal fatigue life N _f1, and judge that can Spacecraft Electronic assembly initial failure device screen out in thermal cycling test, whether the circulating cycle issue N of thermal cycling test alternatives is greater than Spacecraft Electronic assembly initial failure device thermal fatigue life N _f1in statistics Spacecraft Electronic assembly, can screen out initial failure device count and initial failure device sum, calculate Spacecraft Electronic component heat cyclic test validity, if thermal cycling test validity is 100%, carry out step 4, if thermal cycling test validity is not 100%, add tight condition, re-start step 2;

Step 4: each components and parts temperature difference Δ of Spacecraft Electronic assembly T under normal running conditions ₂calculate: according to the requirement of Spacecraft Electronic component design, determine Spacecraft Electronic assembly normal working temperature section, the Spacecraft Electronic assembly digital prototype that utilizes business software and step 2 to set up, calculates each components and parts temperature difference Δ of Spacecraft Electronic assembly T under normal running conditions ₂;

Step 5: Spacecraft Electronic component heat cyclic test breakdown diagnosis: determine Spacecraft Electronic assembly proper device solder joint useful area S ₂, according to the poor Δ T of each device temperature of Spacecraft Electronic assembly under step 2 thermal cycling test condition ₁, calculate Spacecraft Electronic assembly proper device solder joint shear strain Δ γ ₂, utilize Coffin-Manson heat fatigue model, calculate Spacecraft Electronic assembly proper device thermal fatigue life N _f2, can obtain the test damage ratio D of thermal cycling test to Spacecraft Electronic assembly proper device ₁;

Step 6: the Spacecraft Electronic assembly breakdown diagnosis of normally working: according to the definite normal solder joint useful area of the Spacecraft Electronic assembly S of step 5 ₂and each components and parts temperature difference Δ of Spacecraft Electronic assembly T under the definite normal running conditions of step 4 ₂, calculate Spacecraft Electronic assembly proper device solder joint shear strain Δ γ ₃, utilize Coffin-Manson heat fatigue model, calculate Spacecraft Electronic assembly proper device thermal fatigue life N _f3, can meet designed life and meet the work damage ratio D of Spacecraft Electronic assembly proper device under normal running conditions ₂;

Step 7: the cyclic test of Spacecraft Electronic component heat is acceptable to be analyzed: the test damage ratio D obtaining according to step 5 ₁the work damage ratio D obtaining with step 6 ₂, the life cycle management damage ratio D of calculating Spacecraft Electronic assembly proper device, judges the life cycle management damage ratio D>1 that whether has proper device in Spacecraft Electronic assembly, if do not exist, carry out step 8, if exist, soften terms, re-start step 2;

Step 8: final definite thermal cycling test scheme is and rational thermal cycling test scheme effective for concrete Spacecraft Electronic assembly.

Wherein, in the thermal cycling test alternatives described in step 1, comprise the maximum temperature T of thermal cycling test ₁, minimum temperature T ₂, maximum temperature immerses time t ₁, minimum temperature immerses time t ₂, temperature cycles periodicity N, rate temperature change v etc.

Thermal cycling test alternatives can require > > or MIL-HDBK-340A < < Test Requirement for Launch with reference to GJB 1027-2005 < < vehicle, Upper Stage, spacecraft testing, Upper-stage, the national and foreign standards such as and Space Vehicles > >, as shown in Table 1.

Wherein, in the Spacecraft Electronic assembly relevant information described in step 2, comprise the relevant information such as device item, packing forms, material, position, size, power consumption of pcb board appearance profile parameter, pcb board layer information, each device of Spacecraft Electronic assembly.The required business software of modeling has PWA, Flotherm etc.The maximum temperature T calculating respectively at thermal cycling test ₁and minimum temperature T ₂in situation, the maximum temperature T ' of each device shell temperature of Spacecraft Electronic assembly ₁and minimum temperature T ' ₂, obtain the poor Δ T of each device temperature of Spacecraft Electronic assembly under thermal cycling test condition ₁, computing formula is as follows

ΔT ₁＝T′ ₁-T′ ₂。

Table one thermal cycling test alternatives

Wherein, at the Spacecraft Electronic assembly initial failure device solder joint useful area S described in step 3 ₁can be obtained by defect solder joint area actual measurement statistics or engineering experience.

Spacecraft Electronic assembly initial failure device solder joint shear strain Δ γ under thermal cycling test condition ₁, computing formula is as follows

${\mathrm{\&Delta;\&gamma;}}_1=0.5\&times;\frac{K_D}{\left(1.38\mathrm{MPa}\right)S_{1}h}\&times;{({\mathrm{\&Delta;\&alpha;LT}}_s-{\mathrm{\&Delta;\&alpha;LT}}_c)}^2$

In formula, Δ γ ₁for initial failure device solder joint shear strain under thermal cycling test condition; K _dfor solder joint bendind rigidity, unit is N/mm; S ₁for initial failure device solder joint useful area, unit is mm ²; H is solder joint height, and unit is mm; Δ α LT _cfor device swell increment, unit is mm, Δ α LT _sfor circuit board swell increment, unit is that mm computing formula is as follows

$\mathrm{\&Delta;\&alpha;}{\mathrm{LT}}_c=\sqrt{{\left(L_x{\mathrm{\&alpha;}}_{\mathrm{cx}}\right)}^2+{\left(L_y{\mathrm{\&alpha;}}_{\mathrm{cy}}\right)}^2}\&times;{\mathrm{\&Delta;T}}_1$

$\mathrm{\&Delta;\&alpha;}{\mathrm{LT}}_s=\sqrt{{\left(L_x{\mathrm{\&alpha;}}_{\mathrm{sx}}\right)}^2+{\left(L_y{\mathrm{\&alpha;}}_{\mathrm{sy}}\right)}^2}\&times;{\mathrm{\&Delta;T}}_1$

In formula, L _xfor solder joint maximum spacing in x direction or pad array span, unit is mm; L _yfor solder joint maximum spacing in y direction or pad array span, unit is mm; α _cxfor the thermal expansivity of device in x direction, unit is ppm/ ℃; α _cyfor the thermal expansivity of device in y direction, unit is ppm/ ℃; α _sxfor the thermal expansivity of circuit board in x direction, unit is ppm/ ℃; α _syfor the thermal expansivity of circuit board in y direction, unit is ppm/ ℃.

Utilize Coffin-Manson heat fatigue model, calculate Spacecraft Electronic assembly initial failure device thermal fatigue life N _f1, computing formula is as follows

$N_{f1}=\frac{1}{2}{\left(\frac{{\mathrm{\&Delta;\&gamma;}}_1}{{2\mathrm{\&epsiv;}}_f^{\&prime;}}\right)}^{\frac{1}{c}}$

In formula, N _f1for initial failure device thermal fatigue life under thermal cycling test condition; Δ γ ₁for initial failure device solder joint shear strain under thermal cycling test condition; ε ' _ffor fatigue ductile coefficient, conventionally get 2 ε ' _f≈ 0.65; C is fatigue ductility index, and computing formula is as follows

$c=-0.442-6\&times;{10}^{-4}{T}_m+1.74\&times;{10}^{-2}\ln (1+\left(\frac{360}{t_D}\right))$

In formula, T _mfor test period solder joint medial temperature, unit is ℃; t _dfor maximum temperature is immersed the time, unit is min.

Can judgement Spacecraft Electronic assembly initial failure device screen out in thermal cycling test, and whether the circulating cycle issue N of thermal cycling test alternatives is greater than Spacecraft Electronic assembly initial failure device thermal fatigue life N _f1, in statistics Spacecraft Electronic assembly, can screen out initial failure device count m and initial failure device sum M, calculate Spacecraft Electronic component heat cyclic test validity E, computing formula is as follows

$E=\frac{m}{M}\&times;100\%$

If thermal cycling test validity is 100%, carry out step 4, if thermal cycling test validity is not 100%, add tight condition, suitably increase maximum temperature T ₁or reduce minimum temperature T ₂or increase temperature cycles periodicity N, re-start step 2.

Wherein, at the Spacecraft Electronic assembly normal working temperature section described in step 4, by Spacecraft Electronic component design, required to determine, mainly comprise the maximum temperature T in normal work ₃, minimum temperature T ₄, maximum temperature immerses time t ₃, minimum temperature immerses time t ₄, rate temperature change v, cycle designed life experience temperature cycles counts N ₁deng.

The Spacecraft Electronic assembly digital prototype that utilizes business software PWA, Flotherm etc. and step 2 to set up, calculates respectively maximum temperature T during normal use ₃and minimum temperature T ₄in situation, the maximum temperature T ' of each device shell temperature of Spacecraft Electronic assembly ₃and minimum temperature T ' ₄, the poor Δ T of each device temperature of spacecraft electronic package in normally being worked ₂, computing formula is as follows

ΔT ₂＝T′ ₃-T′ ₄

Wherein, at the Spacecraft Electronic assembly proper device solder joint useful area S described in step 5 ₂can be obtained by normal solder joint useful area actual measurement statistics or device handbook.

Spacecraft Electronic assembly proper device solder joint shear strain Δ γ under thermal cycling test condition ₂, computing formula is as follows

${\mathrm{\&Delta;\&gamma;}}_2=0.5\&times;\frac{K_D}{\left(1.38\mathrm{MPa}\right)S_{2}h}\&times;{({\mathrm{\&Delta;\&alpha;LT}}_s-{\mathrm{\&Delta;\&alpha;LT}}_c)}^2$

In formula, Δ γ ₂for proper device solder joint shear strain under thermal cycling test condition; K _dfor solder joint bendind rigidity, unit is N/mm; S ₂for proper device solder joint useful area, unit is mm ²; H is solder joint height, and unit is mm; Δ α LT _cfor device swell increment, Δ α LT _sfor circuit board swell increment, computing formula is as follows

$\mathrm{\&Delta;\&alpha;}{\mathrm{LT}}_c=\sqrt{{\left(L_x{\mathrm{\&alpha;}}_{\mathrm{cx}}\right)}^2+{\left(L_y{\mathrm{\&alpha;}}_{\mathrm{cy}}\right)}^2}\&times;{\mathrm{\&Delta;T}}_1$

$\mathrm{\&Delta;\&alpha;}{\mathrm{LT}}_s=\sqrt{{\left(L_x{\mathrm{\&alpha;}}_{\mathrm{sx}}\right)}^2+{\left(L_y{\mathrm{\&alpha;}}_{\mathrm{sy}}\right)}^2}\&times;{\mathrm{\&Delta;T}}_1$

In formula, L _xfor solder joint maximum spacing in x direction or pad array span, unit is in; L _yfor solder joint maximum spacing in y direction or pad array span, unit is in; α _cxfor the thermal expansivity of device in x direction, unit is ppm/ ℃; α _cyfor the thermal expansivity of device in y direction, unit is ppm/ ℃; α _sxfor the thermal expansivity of circuit board in x direction, unit is ppm/ ℃; α _syfor the thermal expansivity of circuit board in y direction, unit is ppm/ ℃.

Utilize Coffin-Manson heat fatigue model, calculate Spacecraft Electronic assembly proper device thermal fatigue life N under thermal cycling test condition _f2, computing formula is as follows

$N_{f2}=\frac{1}{2}{\left(\frac{{\mathrm{\&Delta;\&gamma;}}_2}{{2\mathrm{\&epsiv;}}_f^{\&prime;}}\right)}^{\frac{1}{c}}$

In formula, N _f2for proper device thermal fatigue life under thermal cycling test condition; Δ γ ₂for proper device solder joint shear strain under thermal cycling test condition; ε ' _ffor fatigue ductile coefficient, conventionally get 2 ε ' _f≈ 0.65; C is fatigue ductility index, and computing formula is as follows

$c=-0.442-6\&times;{10}^{-4}{T}_m+1.74\&times;{10}^{-2}\ln (1+\left(\frac{360}{t_D}\right))$

In formula, T _mfor test period solder joint medial temperature, unit is ℃; t _dfor maximum temperature is immersed the time, unit is min.

Calculate the test damage ratio D of thermal cycling test to Spacecraft Electronic assembly proper device ₁, computing formula is as follows

$D_1=\frac{N}{N_{f2}}$

Wherein, Spacecraft Electronic assembly proper device solder joint shear strain Δ γ under the spacecraft normal running conditions described in step 6 ₃, computing formula is as follows

${\mathrm{\&Delta;\&gamma;}}_3=0.5\&times;\frac{K_D}{\left(1.38\mathrm{MPa}\right)S_{2}h}\&times;{({\mathrm{\&Delta;\&alpha;LT}}_s-{\mathrm{\&Delta;\&alpha;LT}}_c)}^2$

In formula, Δ γ ₃for proper device solder joint shear strain under normal running conditions; K _dfor solder joint bendind rigidity, unit is N/mm; S ₂for proper device solder joint useful area, unit is mm ²; H is solder joint height, and unit is mm; Δ α LT _cfor device swell increment, Δ α LT _sfor circuit board swell increment, computing formula is as follows

$\mathrm{\&Delta;\&alpha;}{\mathrm{LT}}_c=\sqrt{{\left(L_x{\mathrm{\&alpha;}}_{\mathrm{cx}}\right)}^2+{\left(L_y{\mathrm{\&alpha;}}_{\mathrm{cy}}\right)}^2}\&times;{\mathrm{\&Delta;T}}_2$

$\mathrm{\&Delta;\&alpha;}{\mathrm{LT}}_s=\sqrt{{\left(L_x{\mathrm{\&alpha;}}_{\mathrm{sx}}\right)}^2+{\left(L_y{\mathrm{\&alpha;}}_{\mathrm{sy}}\right)}^2}\&times;{\mathrm{\&Delta;T}}_2$

In formula, L _xfor solder joint maximum spacing in x direction or pad array span, unit is in; L _yfor solder joint maximum spacing in y direction or pad array span, unit is in; α _cxfor the thermal expansivity of device in x direction, unit is ppm/ ℃; α _cyfor the thermal expansivity of device in y direction, unit is ppm/ ℃; α _sxfor the thermal expansivity of circuit board in x direction, unit is ppm/ ℃; α _syfor the thermal expansivity of circuit board in y direction, unit is ppm/ ℃.

Utilize Coffin-Manson heat fatigue model, calculate Spacecraft Electronic assembly proper device thermal fatigue life N under normal running conditions _f3, computing formula is as follows

$N_{f3}=\frac{1}{2}{\left(\frac{{\mathrm{\&Delta;\&gamma;}}_3}{{2\mathrm{\&epsiv;}}_f^{\&prime;}}\right)}^{\frac{1}{c}}$

In formula, N _f3for proper device thermal fatigue life under normal running conditions; Δ γ ₃for proper device solder joint shear strain under normal running conditions; ε ' _ffor fatigue ductile coefficient, conventionally get 2 ε ' _f≈ 0.65; C is fatigue ductility index, and computing formula is as follows

$c=-0.442-6\&times;{10}^{-4}{T}_m+1.74\&times;{10}^{-2}\ln (1+\left(\frac{360}{t_D}\right))$

In formula, T _mfor solder joint medial temperature under normal running conditions, unit is ℃; t _dfor maximum temperature under normal running conditions is immersed the time, unit is min.

Calculate the test damage ratio D normally working to Spacecraft Electronic assembly proper device ₂, computing formula is as follows

$D_2=\frac{N_1}{N_{f3}}$

Wherein, at the aerospace electron assembly proper device life cycle management damage ratio D described in step 7, computing formula is as follows

D＝D ₁+D ₂

Judge the life cycle management damage ratio D>1 that whether has proper device in Spacecraft Electronic assembly, if do not exist, carry out step 8, if exist, soften terms, suitably reduce maximum temperature T ₁or increase minimum temperature T ₂or reduce temperature cycles periodicity N, re-start step 2.

Wherein, in step 8, final definite thermal cycling test scheme is to meet thermal cycling test validity and the acceptable scheme requiring of thermal cycling test, and final Output rusults is the maximum temperature T of thermal cycling test scheme ₁, minimum temperature T ₂, maximum temperature immerses time t ₁, minimum temperature immerses time t ₂, temperature cycles periodicity N, rate temperature change v.

The invention provides a kind of Spacecraft Electronic component heat cyclic test scheme and determine method, its advantage mainly contains: the present invention is directed to concrete Spacecraft Electronic assembly, analyze thermal cycling test acceptable to the thermal cycling test validity of Spacecraft Electronic assembly and thermal cycling test, have more specific aim and rationality.The present invention is based on Coffin-Manson thermal fatigue life theory and Miner linear cumulative damage law, by computational analysis, draw effective and rational thermal cycling test scheme, have workable, save the advantages such as experimentation cost, can facilitate and promptly formulate aerospace electron product thermal cycling test scheme.

(4) accompanying drawing explanation:

Fig. 1 is the implementing procedure that Spacecraft Electronic component heat cyclic test scheme is determined method

Number in the figure and symbol description are as follows:

T ₁maximum temperature for thermal cycling test

T ₂minimum temperature for thermal cycling test

T ₁for the maximum temperature of thermal cycling test is immersed the time

T ₂for the minimum temperature of thermal cycling test is immersed the time

N is the temperature cycles periodicity of thermal cycling test

V is the rate temperature change of thermal cycling test

T ' ₁maximum temperature for each device shell temperature of Spacecraft Electronic assembly under thermal cycling test condition

T ' ₂minimum temperature for each device shell temperature of Spacecraft Electronic assembly under thermal cycling test condition

Δ T ₁for each device temperature of Spacecraft Electronic assembly under thermal cycling test condition poor

S ₁for Spacecraft Electronic assembly initial failure device solder joint useful area

Δ γ ₁for Spacecraft Electronic assembly initial failure device solder joint shear strain under thermal cycling test condition

K _dfor solder joint bendind rigidity

H is solder joint height

Δ α LT _cfor device swell increment

Δ α LT _sfor circuit board swell increment

L _xfor solder joint maximum spacing in x direction or pad array span

L _yfor solder joint maximum spacing in y direction or pad array span

α _cxthermal expansivity for device in x direction

α _cythermal expansivity for device in y direction

α _sxthermal expansivity for circuit board in x direction

α _sythermal expansivity for circuit board in y direction

N _f1for Spacecraft Electronic assembly initial failure device thermal fatigue life

Δ γ ₁for initial failure device solder joint shear strain under thermal cycling test condition

ε ' _ffor fatigue ductile coefficient

T _mfor test period solder joint medial temperature

T _dfor maximum temperature is immersed the time

M can screen out initial failure device count in Spacecraft Electronic assembly

M is initial failure device sum in Spacecraft Electronic assembly

E is Spacecraft Electronic component heat cyclic test validity

T ₃for the maximum temperature in normal work

T ₄for the minimum temperature in normal work

T ₃for the maximum temperature in normal work is immersed the time

T ₄for the minimum temperature in normal work is immersed the time

V is the rate temperature change in normal work

N ₁for cycle designed life experience temperature cycles number

T ' ₃maximum temperature for each device shell temperature of spacecraft electronic package in normal work

T ' ₄minimum temperature for each device shell temperature of spacecraft electronic package in normal work

Δ T ₂for each device temperature of spacecraft electronic package in normal work poor

S ₂for Spacecraft Electronic assembly proper device solder joint useful area

Δ γ ₂for Spacecraft Electronic assembly proper device solder joint shear strain under thermal cycling test condition

S ₂for proper device solder joint useful area

N _f2for Spacecraft Electronic assembly proper device thermal fatigue life under thermal cycling test condition

D ₁test damage ratio for Spacecraft Electronic assembly proper device

Δ γ ₃for spacecraft electronic package proper device solder joint shear strain in normal work

N _f3for spacecraft electronic package proper device thermal fatigue life in normal work

D ₂work damage ratio for spacecraft electronic package proper device in normal work

D is aerospace electron assembly proper device life cycle management damage ratio

(5) embodiment:

Below in conjunction with concrete case study on implementation, the Spacecraft Electronic component heat cyclic test scheme for concrete Spacecraft Electronic assembly of the present invention is determined to method is elaborated.

Case: the rail control engine parameter collection plate that is applied to spacecraft

It is example that the rail control engine parameter collection plate that is applied to spacecraft is take in the present invention, illustrates for the Spacecraft Electronic component heat cyclic test scheme of concrete Spacecraft Electronic assembly and determines method.

Rail control engine parameter collection plate information and collection plate application requirements are as follows:

(1) rail control engine parameter collection plate: this parameter acquisition plate is for certain type synchronous orbit communications satellite, the propelling module that is positioned at satellite, is of a size of 96mm * 86mm * 1.53mm, and total power consumption is 1.4W, comprise 11 electronic devices and components of 8 class, device inventory as shown in Table 4.

(2) application requirements: this synchronous orbit communications satellite designed life is 15 years, and the temperature cycles cycle is 24 hours, and wherein, the warm change time is 6 hours, and it is 6 hours that temperature is immersed the time.Generally, the environment temperature of propelling module is-10 to+40 ℃, and during collection plate work, the application requirements of temperature environment as shown in Table 2.

Table two collection plate operating ambient temperature correlation parameter

According to the flow process of Fig. 1, the method concrete steps are as follows:

Step 1: determine thermal cycling test alternatives, tentatively determine the maximum temperature T of thermal cycling test ₁be 60 ℃, minimum temperature T ₂for-25 ℃, maximum temperature is immersed time t ₁for 60min, minimum temperature is immersed time 60min, temperature cycles periodicity N is 25, and rate temperature change v is 4 ℃/min.

Step 2: collect Spacecraft Electronic assembly relevant information, pcb board is of a size of 96mm * 86mm * 1.53mm, the flaggy information of pcb board as shown in Table 3, the relevant information such as the device item of each device, packing forms, material, position, size, power consumption in Spacecraft Electronic assembly, as shown in Table 4.

The flaggy information of table three pcb board

Flaggy	Cover sheet materials	Flaggy thickness (mm)	Flaggy metal material	Flaggy tenor (%)
					Top layer	epoxyF	0.06	Cu	23.17
Ground floor	FR4	0.43	Cu	2.56
					Power	epoxyF	0.06	Cu	95.64
The second layer	FR4	0.43	Cu	2.56
					GND	epoxyF	0.06	Cu	96.12
The 3rd layer	FR4	0.43	Cu	2.56
					Bottom	epoxyF	0.06	Cu	11.23

Table four rail control engine parameter collection plate device inventory

Utilize the business softwares such as PWA, Flotherm, the maximum temperature T calculating respectively at thermal cycling test ₁and minimum temperature T ₂the maximum temperature T ' of each device shell temperature of Spacecraft Electronic assembly ₁and minimum temperature T ' ₂, obtain the poor Δ T of each device temperature of Spacecraft Electronic assembly under thermal cycling test condition ₁, as shown in Table 5.

Under table five thermal cycling test condition, each device temperature of Spacecraft Electronic assembly is poor

Step 3: collect correlation parameter, calculate each device and circuit board swell increment thereof, as shown in Table 6.Utilize correlation formula and Spacecraft Electronic assembly initial failure correlation parameter, calculate Spacecraft Electronic assembly initial failure device solder joint shear strain Δ γ under thermal cycling test condition ₁, as shown in Table 7.Utilize Coffin-Manson heat fatigue model, calculate aerospace electron assembly initial failure device thermal fatigue life N _f1, as shown in Table 8.

Each device of Spacecraft Electronic assembly and circuit board swell increment thereof under table six thermal cycling test condition

Spacecraft Electronic assembly initial failure device solder joint shear strain under table seven thermal cycling test condition

Numbering	Item	ΔT 1(℃)	ΔαLT c	ΔαLT s	K D	S 1	h	Δγ 1
									1	C101	84.9656	0.0098	0.0126	540000000	0.306	0.2	0.1965
2	C102	84.9493	0.0098	0.0126	35600000	0.0306	0.2	0.1295
									3	C103	84.93637	0.0098	0.0126	35600000	0.0306	0.2	0.1295
4	D101	84.89719	0.0216	0.0611	1970	0.025893	0.25	0.2684
									5	D102	84.87448	0.0208	0.0587	2400	0.025893	0.25	0.2772
6	D103	84.95754	0.0041	0.0117	34000	0.00135	0.2	0.1522
									7	D104	84.96317	0.0100	0.0117	13300000	0.00135	0.2	0.1491
8	D105	84.96311	0.0136	0.0158	10700000	0.003906	0.145	0.1603
									9	D106	84.9197	0.0098	0.0276	5800	0.0045	0.25	0.1875
10	D107	84.94361	0.0064	0.0181	106000	0.01593	0.25	0.1807

?

Numbering	Item	ΔT 1(℃)	ΔαLT c	ΔαLT s	K D	S 1	h	Δγ 1
									11	D108	84.90089	0.0208	0.0587	2300	0.025893	0.25	0.2656

Table eight thermal cycling test efficiency analysis

Numbering	Item	ΔT 1(℃)	T m	t D	c	2ε' f	Δγ 1	N f1	＜N
										1	C101	84.9656	17.8157	60	-0.418831	0.65	0.1965	8.7	Be
2	C102	84.9493	17.96507	60	-0.41892	0.65	0.1295	23.5	Be
										3	C103	84.93637	18.0966	60	-0.418999	0.65	0.1295	23.5	Be
4	D101	84.89719	18.74909	60	-0.419391	0.65	0.2684	4.1	Be
										5	D102	84.87448	18.98552	60	-0.419532	0.65	0.2772	3.8	Be
6	D103	84.95754	18.17347	60	-0.419045	0.65	0.1522	16.0	Be
										7	D104	84.96317	18.11839	60	-0.419012	0.65	0.1491	16.8	Be
8	D105	84.96311	18.13628	60	-0.419023	0.65	0.1603	14.1	Be
										9	D106	84.9197	18.79857	60	-0.41942	0.65	0.1875	9.7	Be
10	D107	84.94361	18.63799	60	-0.419324	0.65	0.1807	10.6	Be
										11	D108	84.90089	18.73037	60	-0.419379	0.65	0.2656	4.2	Be

As shown in Table 8, in Spacecraft Electronic assembly, can screen out initial failure device count m is 11, and initial failure device sum M is 11, and Spacecraft Electronic component heat cyclic test validity E is 100%.

Due to Spacecraft Electronic component heat cyclic test validity, E is 100%, carry out step 4.

Step 4: Spacecraft Electronic assembly normal working temperature section, by Spacecraft Electronic component design, required to determine, mainly comprise the maximum temperature T in normal work ₃be 50 ℃, minimum temperature T ₄for-10 ℃, maximum temperature is immersed time t ₃be 6 hours, minimum temperature is immersed time t ₄be 6 hours, rate temperature change v is 10 ℃/h, and cycle designed life experience temperature cycles is counted N ₁it is 5475 times.

Utilize the business softwares such as PWA, Flotherm, calculate respectively maximum temperature T in normal work ₃and minimum temperature T ₄the maximum temperature T ' of each device shell temperature of Spacecraft Electronic assembly ₃and minimum temperature T ' ₄, the poor Δ T of each device temperature of spacecraft electronic package in normally being worked ₂, as shown in Table 9.

In the normal work of table nine, each device temperature of spacecraft electronic package is poor

Step 5: collect correlation parameter, calculate each device and circuit board swell increment thereof, as shown in Table 6.Utilize correlation formula and spacecraft normal electrical sub-component correlation parameter, calculate Spacecraft Electronic assembly proper device solder joint shear strain Δ γ under thermal cycling test condition ₂, as shown in Table 10.Utilize Coffin-Manson heat fatigue model, calculate aerospace electron assembly proper device thermal fatigue life N under thermal cycling test condition _f2, and calculate the test damage ratio D of thermal cycling test to Spacecraft Electronic assembly proper device ₁, as shown in Table 8.

Spacecraft Electronic assembly initial failure device solder joint shear strain under table ten thermal cycling test condition

Numbering	Item	ΔT 1(℃)	ΔαLT c	ΔαLT s	K D	S 2	h	Δγ 2
									1	C101	84.9656	0.0098	0.0126	540000000	1.02	0.2	0.0590
2	C102	84.9493	0.0098	0.0126	35600000	1.02	0.2	0.0039
									3	C103	84.93637	0.0098	0.0126	35600000	1.02	0.2	0.0039
4	D101	84.89719	0.0216	0.0611	1970	0.3699	0.25	0.0188
									5	D102	84.87448	0.0208	0.0587	2400	0.3699	0.25	0.0194
6	D103	84.95754	0.0041	0.0117	34000	0.09	0.2	0.0023
									7	D104	84.96317	0.01	0.0117	13300000	0.09	0.2	0.0022
8	D105	84.96311	0.0136	0.0158	10700000	0.1302	0.145	0.0048
									9	D106	84.9197	0.0098	0.0276	5800	0.09	0.25	0.0094
10	D107	84.94361	0.0038	0.0106	106000	0.531	0.25	0.0054
									11	D108	84.90089	0.0122	0.0346	2300	0.3699	0.25	0.0186

The test damage ratio of table ten thermal cycling test to Spacecraft Electronic assembly proper device

Numbering	Item	ΔT 1(℃)	T m	t D	c	2ε' f	Δγ 2	N f2	D 1
										1	C101	84.9656	17.8157	60	-0.418831	0.65	0.0590	154.7	0.1616
2	C102	84.9493	17.96507	60	-0.41892	0.65	0.0039	101185.2	0.0002
										3	C103	84.93637	18.0966	60	-0.418999	0.65	0.0039	101014.5	0.0002
4	D101	84.89719	18.74909	60	-0.419391	0.65	0.0188	2340.6	0.0107
										5	D102	84.87448	18.98552	60	-0.419532	0.65	0.0194	2335.4	0.0107
6	D103	84.95754	18.17347	60	-0.419045	0.65	0.0023	375345.5	0.0001
										7	D104	84.96317	18.11839	60	-0.419012	0.65	0.0022	375746.2	0.0001
8	D105	84.96311	18.13628	60	-0.419023	0.65	0.0048	61094.0	0.0004
										9	D106	84.9197	18.79857	60	-0.41942	0.65	0.0094	12289.9	0.0020
10	D107	84.94361	18.63799	60	-0.419324	0.65	0.0054	44872.4	0.0006
										11	D108	84.90089	18.73037	60	-0.419379	0.65	0.0186	2340.8	0.0107

Step 6: according to each components and parts temperature difference Δ of Spacecraft Electronic assembly T under the normal phase of solder joint related parameter of Spacecraft Electronic assembly and normal running conditions ₂, calculate Spacecraft Electronic assembly device thermal expansion amount and proper device solder joint shear strain Δ γ under normal running conditions ₃, as shown in table ten two and table ten three.Utilize Coffin-Manson heat fatigue model, calculate Spacecraft Electronic assembly proper device thermal fatigue life N under normal running conditions _f3, and calculate the work damage ratio D of Spacecraft Electronic assembly proper device under normal running conditions ₂, as shown in table ten four.

Spacecraft Electronic assembly device thermal expansion amount under table ten two normal running conditions

Spacecraft Electronic assembly proper device solder joint shear strain under table ten three normal running conditions

Numbering	Item	ΔT 1(℃)	ΔαLT c	ΔαLT s	K D	S 2	h	Δγ 3
									1	C101	50.01978	0.0057	0.0074	540000000	1.02	0.2	0.0075
2	C102	50.01195	0.0057	0.0074	35600000	1.02	0.2	0.0005
									3	C103	50.00648	0.0057	0.0074	35600000	1.02	0.2	0.0005
4	D101	50.06008	0.0128	0.0360	1970	0.3699	0.25	0.0023
									5	D102	50.01005	0.0123	0.0346	2400	0.3699	0.25	0.0023
6	D103	49.76564	0.0024	0.0068	34000	0.09	0.2	0.0003
									7	D104	49.76512	0.0059	0.0068	13300000	0.09	0.2	0.0002
8	D105	51.02652	0.0082	0.0095	10700000	0.1302	0.145	0.0007
									9	D106	50.16632	0.0058	0.0163	5800	0.09	0.25	0.0012
10	D107	49.86146	0.0038	0.0106	106000	0.531	0.25	0.0006
									11	D108	49.98466	0.0122	0.0346	2300	0.3699	0.25	0.0022

The work damage ratio of Spacecraft Electronic assembly proper device under table ten four normal running conditions

Numbering	Item	ΔT 1(℃)	T m	t D	c	2ε' f	Δγ 3	N f3	D 2
										1	C101	50.01978	15.34074	60	-0.417346	0.65	0.0075	22304.9	0.2455
2	C102	50.01195	15.493395	60	-0.417437	0.65	0.0005	14935709.4	0.0004
										3	C103	50.00648	15.6279	60	-0.417518	0.65	0.0005	14910884.6	0.0004
4	D101	50.06008	16.32446	60	-0.417936	0.65	0.0023	382475.0	0.0143
										5	D102	50.01005	16.545985	60	-0.418069	0.65	0.0023	379769.5	0.0144
6	D103	49.76564	15.57503	60	-0.417486	0.65	0.0003	72000112.4	0.0001
										7	D104	49.76512	15.51722	60	-0.417451	0.65	0.0002	80130252.0	0.0001
8	D105	51.02652	16.16586	60	-0.417841	0.65	0.0007	6366164.9	0.0009
										9	D106	50.16632	16.41722	60	-0.417991	0.65	0.0012	1921589.0	0.0028
10	D107	49.86146	16.09365	60	-0.417797	0.65	0.0006	7691387.6	0.0007
										11	D108	49.98466	16.26635	60	-0.417901	0.65	0.0022	383406.9	0.0143

Step 7: the cyclic test of Spacecraft Electronic component heat is acceptable to be analyzed: the test damage ratio D obtaining according to step 5 ₁the work damage ratio D obtaining with step 6 ₂, calculate the life cycle management damage ratio D of Spacecraft Electronic assembly proper device, whether the life cycle management damage ratio D that judges proper device in Spacecraft Electronic assembly >1, as shown in table ten five.Owing to there not being the life cycle management damage ratio D>1 of proper device in Spacecraft Electronic assembly, carry out step 8.

The life cycle management damage ratio D of table ten five Spacecraft Electronic assembly proper device

Numbering	Item	D 1	D 2	D	> 1 whether
						1	C101	0.1616	0.2455	0.4071	No
2	C102	0.0002	0.0004	0.0006	No
						3	C103	0.0002	0.0004	0.0006	No
4	D101	0.0107	0.0143	0.0250	No
						5	D102	0.0107	0.0144	0.0251	No
6	D103	0.0001	0.0001	0.0001	No
						7	D104	0.0001	0.0001	0.0001	No
8	D105	0.0004	0.0009	0.0013	No
						9	D106	0.0020	0.0028	0.0049	No
10	D107	0.0006	0.0007	0.0013	No
						11	D108	0.0107	0.0143	0.0250	No

Step 8: final this thermal cycling test scheme of determining is and rational thermal cycling test scheme effective for parameter acquisition plate, as shown in table ten six.

Table ten six thermal cycling test schemes

Claims (9)

1. method is determined in a Spacecraft Electronic component heat cyclic test program analysis, it is characterized in that: utilize heat analysis, Coffin-Manson model and life cycle theory, acceptable based on thermal cycling test validity and thermal cycling test, analyze thermal cycling test scheme to specifying validity and the damage of Spacecraft Electronic assembly, according to this method flow process, iteration judgement, finally determines suitable Spacecraft Electronic component heat cyclic test scheme, and the method concrete steps are as follows:

Step 1: the determining of thermal cycling test alternatives;

Step 2: each components and parts temperature difference Δ of Spacecraft Electronic assembly T under thermal cycling test condition ₁calculate;

Step 3: Spacecraft Electronic component heat cyclic test efficiency analysis, if thermal cycling test validity is 100%, carry out step 4, if thermal cycling test validity is not 100%, add tight condition, re-start step 2;

Step 4: each components and parts temperature difference Δ of Spacecraft Electronic assembly T under normal running conditions ₂calculate;

Step 5: Spacecraft Electronic component heat cyclic test breakdown diagnosis;

Step 6: the Spacecraft Electronic assembly breakdown diagnosis of normally working;

Step 7: the cyclic test of Spacecraft Electronic component heat is acceptable to be analyzed, and judges the life cycle management damage ratio D>1 that whether has proper device in Spacecraft Electronic assembly, if do not exist, carry out step 8, if exist, soften terms, re-start step 2;

Step 8: final definite thermal cycling test scheme is and rational thermal cycling test scheme effective for concrete Spacecraft Electronic assembly.

2. determining of thermal cycling test alternatives according to claim 1, is characterized in that: at the main of the scheme of thermal cycling test described in step 2, determine that parameter is maximum temperature T ₁, minimum temperature T ₂, maximum temperature immerses time t ₁, minimum temperature immerses time t ₂, temperature cycles periodicity N, rate temperature change v.

3. each components and parts temperature difference Δ of Spacecraft Electronic assembly T under thermal cycling test condition according to claim 1 ₁calculate, it is characterized in that: described in step 2, utilizing related software, calculating the shell temperature of each components and parts of aerospace electron product component under each temperature levels of thermal cycling test, and drawing the shell temperature difference T of each components and parts of aerospace electron product component ₁.

4. Spacecraft Electronic component heat cyclic test efficiency analysis according to claim 1, is characterized in that:

(1) at the assembly of Spacecraft Electronic described in step 3 initial failure device solder joint useful area S ₁can be obtained by defect solder joint area actual measurement statistics or engineering experience;

(2) whole injections that Spacecraft Electronic assembly initial failure is device fault under the condition of thermal cycling test described in step 3;

(3) Spacecraft Electronic assembly initial failure device solder joint shear strain Δ γ under the condition of thermal cycling test described in step 3 ₁computing formula

${\mathrm{\&Delta;\&gamma;}}_1=0.5\&times;\frac{K_D}{\left(1.38\mathrm{MPa}\right)S_{1}h}\&times;{({\mathrm{\&Delta;\&alpha;LT}}_s-{\mathrm{\&Delta;\&alpha;LT}}_c)}^2$

In formula, Δ γ ₁for initial failure device solder joint shear strain under thermal cycling test condition, K _dfor solder joint bendind rigidity, unit is N/mm, S ₁for initial failure device solder joint useful area, unit is mm ², h is solder joint height, unit is mm;

(4) in the component heat of Spacecraft Electronic described in step 3 cyclic test validity E computing formula

$E=\frac{m}{M}\&times;100\%$

In formula, E is Spacecraft Electronic component heat cyclic test validity, and m can screen out initial failure device count in Spacecraft Electronic assembly, and M is initial failure device sum in Spacecraft Electronic assembly;

(5) if be 100% in the validity of thermal cycling test described in step 3, carry out step 4, if thermal cycling test validity is not 100%, add tight condition, suitably increase maximum temperature T ₁or reduce minimum temperature T ₂or increase temperature cycles periodicity N, re-start step 2.

5. each components and parts temperature difference Δ of Spacecraft Electronic assembly T under normal running conditions according to claim 1 ₂, it is characterized in that: described in step 4, utilizing related software, calculating the shell temperature of each components and parts of aerospace electron product component under normal each temperature levels of work, and drawing the shell temperature difference T of each components and parts of aerospace electron product component ₂.

6. Spacecraft Electronic component heat cyclic test breakdown diagnosis according to claim 1, is characterized in that:

(1) at the assembly of Spacecraft Electronic described in step 5 proper device solder joint useful area S ₂can be obtained by normal solder joint useful area actual measurement statistics or device handbook;

(2) Spacecraft Electronic assembly proper device solder joint shear strain Δ γ under the condition of thermal cycling test described in step 5 ₂computing formula

${\mathrm{\&Delta;\&gamma;}}_2=0.5\&times;\frac{K_D}{\left(1.38\mathrm{MPa}\right)S_{2}h}\&times;{({\mathrm{\&Delta;\&alpha;LT}}_s-{\mathrm{\&Delta;\&alpha;LT}}_c)}^2$

In formula, Δ γ ₂for initial failure device solder joint shear strain under thermal cycling test condition, K _dfor solder joint bendind rigidity, unit is N/mm, S ₂for proper device solder joint useful area, unit is mm ², h is solder joint height, unit is mm;

(3) the test damage ratio D to Spacecraft Electronic assembly proper device at thermal cycling test described in step 5 ₁computing formula

$D_1=\frac{N}{N_{f2}}$

In formula, D ₁for the test damage ratio of thermal cycling test to Spacecraft Electronic assembly proper device, the temperature cycles periodicity that N is thermal cycling test, N _f2for proper device thermal fatigue life under thermal cycling test condition.

7. the Spacecraft Electronic assembly according to claim 1 breakdown diagnosis of normally working, is characterized in that:

(1) Spacecraft Electronic assembly proper device solder joint shear strain Δ γ under the normal running conditions of spacecraft described in step 6 ₃computing formula

${\mathrm{\&Delta;\&gamma;}}_3=0.5\&times;\frac{K_D}{\left(1.38\mathrm{MPa}\right)S_{2}h}\&times;{({\mathrm{\&Delta;\&alpha;LT}}_s-{\mathrm{\&Delta;\&alpha;LT}}_c)}^2$

In formula, Δ γ ₃for initial failure device solder joint shear strain under thermal cycling test condition, K _dfor solder joint bendind rigidity, unit is N/mm, S ₂for proper device solder joint useful area, unit is mm ², h is solder joint height, unit is mm;

(2) at the test damage ratio D normally working to Spacecraft Electronic assembly proper device described in step 6 ₂computing formula

$D_2=\frac{N_1}{N_{f3}}$

In formula, D ₂for the test damage ratio of normally working to Spacecraft Electronic assembly proper device, N ₁for cycle designed life experience temperature cycles number, N _f3for proper device thermal fatigue life under normal running conditions.

8. Spacecraft Electronic component heat cyclic test according to claim 1 is acceptable analyzes, and it is characterized in that:

(1) in the life cycle management damage ratio D of the assembly of Spacecraft Electronic described in step 7 proper device computing method

D＝D ₁+D ₂

In formula, D is the life cycle management damage ratio of Spacecraft Electronic assembly proper device, D ₁for the test damage ratio of thermal cycling test to Spacecraft Electronic assembly proper device, D ₂for the test damage ratio of normally working to Spacecraft Electronic assembly proper device;

(2) in judging Spacecraft Electronic assembly described in step 7, whether there is the life cycle management damage ratio D>1 of proper device, if do not exist, carry out step 8, if exist, soften terms, suitably reduce maximum temperature T ₁or increase minimum temperature T ₂or reduce temperature cycles periodicity N, re-start step 2.

9. final definite thermal cycling test scheme according to claim 1, is characterized in that: at final Output rusults described in step 8, be the maximum temperature T of thermal cycling test scheme ₁, minimum temperature T ₂, maximum temperature immerses time t ₁, minimum temperature immerses time t ₂, temperature cycles periodicity N, rate temperature change v.