CN104007351B - A kind of Spacecraft Electronic component heat cyclic test scheme determines method - Google Patents

Abstract

A kind of Spacecraft Electronic component heat cyclic test scheme determines that method comprises the steps: step 1: the determination of thermal cycling test alternative；Step 2: Spacecraft Electronic assembly each components and parts temperature difference Δ T under the conditions of thermal cycling test₁Calculate；Step 3: Spacecraft Electronic component heat cyclic test efficiency analysis；Step 4: Spacecraft Electronic assembly each components and parts temperature difference Δ T under normal running conditions₂Calculate；Step 5: Spacecraft Electronic component heat cyclic test breakdown diagnosis；Step 6: Spacecraft Electronic assembly normally works breakdown diagnosis；Step 7: Spacecraft Electronic component heat cyclic test acceptability is analyzed；Step 8: the thermal cycling test scheme finally determined.The method can avoid thermal cycling test scheme that Spacecraft Electronic assembly causes overtesting and undertesting, reduces Spacecraft Electronic assembly application risk in space technology.

Description

A kind of Spacecraft Electronic component heat cyclic test scheme determines method

(1) technical field:

The present invention relates to a kind of Spacecraft Electronic component heat cyclic test scheme and determine method, it is adaptable to evaluate different space flight Device electronic building brick is used the scheme of various thermal cycling test scheme the most effectively, is devoted to avoid thermal cycling test scheme Spacecraft Electronic assembly is caused overtesting and undertesting, reduces Spacecraft Electronic assembly application risk in space technology, Belong to spacecraft thermal cycling test technical field.

(2) background technology:

In Spacecraft Electronic product component development process, the most effective thermal cycling test, Spacecraft Electronic can be screened out The potential initial failure of assembly, it is ensured that the reliability of Spacecraft Electronic assembly and service life.

At present in engineering reality, Spacecraft Electronic component heat cyclic test Main Basis GJB 1027-2005 " vehicle, Upper Stage, spacecraft testing require " and carry out.But GJB 1027-2005 adopts for different Spacecraft Electronic product components With identical test requirements document, the defect characteristics of concrete Spacecraft Electronic product component and working section are lacked effectively for Property.Practical Project also constantly occurs broken down within cycle projected life by the Spacecraft Electronic assembly of thermal cycling test Case.

On the other hand, generally analyze thermal cycling test scheme, only consider thermal cycling test scheme effectiveness, i.e. sudden and violent in test The number of faults of dew and the ratio of total failare number, do not consider the unnecessary damage that product causes due to test overstress.Analyze really Determine thermal cycling test scheme, thermal cycling test must be considered to the undertesting of Spacecraft Electronic assembly and overtesting risk, with Obtain reasonable and effective thermal cycling test scheme.

(3) summary of the invention:

1, purpose: the present invention seeks to: provide a kind of Spacecraft Electronic component heat cyclic test scheme to determine method, for Concrete Spacecraft Electronic assembly, considers undertesting and overtesting, to determine the most effective thermal cycling test scheme.

2, technical scheme: the present invention is with thermal cycling test alternative as object of study, for concrete Spacecraft Electronic group Part is calculated by electronic building brick each device temperature difference and calculates with solder joint plasticity shear strain, utilizes the Coffin-Manson heat exhaustion longevity Fate opinion and Miner linear cumulative damage law, divide with thermal cycling test acceptability based on thermal cycling test efficiency analysis Analysis, sets up Spacecraft Electronic component heat cyclic test scheme and determines method, and it comprises the steps:

Step 1: the determination of thermal cycling test alternative: design requirement according to Spacecraft Electronic assembly, primarily determine that heat Cyclic test alternative；

Step 2: Spacecraft Electronic assembly each components and parts temperature difference Δ T under the conditions of thermal cycling test₁Calculate: collect spacecraft Electronic building brick relevant information, utilizes business software, sets up Spacecraft Electronic assembly digital prototype, and the heat determined according to step 1 is followed Ring test alternative, calculates Spacecraft Electronic assembly each device temperature difference Δ T under the conditions of thermal cycling test₁；

Step 3: Spacecraft Electronic component heat cyclic test efficiency analysis: determine Spacecraft Electronic assembly initial failure device Part solder joint effective area S₁, according to Spacecraft Electronic assembly each device temperature difference Δ T under the conditions of step 2 thermal cycling test₁, calculate Spacecraft Electronic assembly initial failure device solder joint shear strain Δ γ₁, utilize Coffin-Manson heat exhaustion model, calculate boat It device electronic building brick initial failure device thermal fatigue life N_f1, and judge that can Spacecraft Electronic assembly initial failure device in heat Screen out in cyclic test, i.e. whether the circulating cycle issue N of thermal cycling test alternative is more than the event in early days of Spacecraft Electronic assembly Barrier device thermal fatigue life N_f1, statistics Spacecraft Electronic assembly can screen out initial failure device count total with initial failure device Number, calculates Spacecraft Electronic component heat cyclic test effectiveness, if thermal cycling test effectiveness is 100%, then carries out step 4, If thermal cycling test effectiveness is not 100%, then adds tight condition, re-start step 2；

Step 4: Spacecraft Electronic assembly each components and parts temperature difference Δ T under normal running conditions₂Calculate: according to spacecraft electricity Sub-component design requirement, determines Spacecraft Electronic assembly normal working temperature section, utilizes the boat that business software and step 2 are set up It device electronic building brick digital prototype, calculates Spacecraft Electronic assembly each components and parts temperature difference Δ T under normal running conditions₂；

Step 5: Spacecraft Electronic component heat cyclic test breakdown diagnosis: determine Spacecraft Electronic assembly proper device solder joint Effective area S₂, according to Spacecraft Electronic assembly each device temperature difference Δ T under the conditions of step 2 thermal cycling test₁, calculate spacecraft Electronic building brick proper device solder joint shear strain Δ γ₂, utilize Coffin-Manson heat exhaustion model, calculate Spacecraft Electronic group Part proper device thermal fatigue life N_f2, the thermal cycling test test damage ratio D to Spacecraft Electronic assembly proper device can be obtained₁；

Step 6: Spacecraft Electronic assembly normally works breakdown diagnosis: the Spacecraft Electronic assembly determined according to step 5 is just Often solder joint effective area S₂And Spacecraft Electronic assembly each components and parts temperature difference Δ T under the normal running conditions that determines of step 4₂, meter Calculate Spacecraft Electronic assembly proper device solder joint shear strain Δ γ₃, utilize Coffin-Manson heat exhaustion model, calculate space flight Device electronic building brick proper device thermal fatigue life N_f3, projected life can be met and meet Spacecraft Electronic under normal running conditions The work damage ratio D of assembly proper device₂；

Step 7: Spacecraft Electronic component heat cyclic test acceptability is analyzed: the test damage ratio obtained according to step 5 D₁The work damage ratio D obtained with step 6₂, calculate the life cycle management damage ratio D of Spacecraft Electronic assembly proper device, it is judged that Whether Spacecraft Electronic assembly existing the life cycle management damage ratio D > 1 of proper device, if not existing, then carrying out step 8, if Exist, then soften terms, re-start step 2；

Step 8: the thermal cycling test scheme finally determined is for for concrete Spacecraft Electronic assembly effectively and reasonably heat Cyclic test scheme.

Wherein, thermal cycling test alternative described in step 1, including the maximum temperature T of thermal cycling test₁, Low temperature T₂, maximum temperature immerse time t₁, minimum temperature immerse time t₂, temperature cycles periodicity N, rate temperature change v Deng.

Thermal cycling test alternative refer to GJB 1027-2005 " vehicle, Upper Stage, spacecraft testing require " or The states such as MIL-HDBK-340A " Test Requirement for Launch, Upper-stage, and Space Vehicles " Inside and outside standard, as shown in Table 1.

Wherein, Spacecraft Electronic device related information described in step 2, including pcb board appearance profile parameter, PCB The relevant letters such as flaggy information, the device item of each device of Spacecraft Electronic assembly, packing forms, material, position, size, power consumption Breath.The required business software of modeling has PWA, Flotherm etc..Calculate the maximum temperature T at thermal cycling test respectively₁And lowest temperature Degree T₂In the case of, the maximum temperature T ' of Spacecraft Electronic assembly each device shell temperature₁And minimum temperature T '₂, obtain thermal cycling test bar Spacecraft Electronic assembly each device temperature difference Δ T under part₁, computing formula is as follows

ΔT₁=T '₁-T′₂。

Table one thermal cycling test alternative

Wherein, described in step 3 Spacecraft Electronic assembly initial failure device solder joint effective area S₁Can be by defect Solder joint area actual measurement statistics or engineering experience obtain.

Spacecraft Electronic assembly initial failure device solder joint shear strain Δ γ under the conditions of thermal cycling test₁, computing formula is such as Under

${\mathrm{\Δ\γ}}_1=0.5\×\frac{K_D}{\left(1.38\mathrm{MPa}\right)S_{1}h}\×{({\mathrm{\Δ\αLT}}_s-{\mathrm{\Δ\αLT}}_c)}^2$

In formula, Δ γ₁For initial failure device solder joint shear strain under the conditions of thermal cycling test；K_DFor solder joint bending rigidity, list Position is N/mm；S₁For initial failure device solder joint effective area, unit is mm²；H is solder joint height, and unit is mm；ΔαLT_cFor Device swell increment, unit is mm, Δ α LT_sFor circuit board swell increment, unit is that mm computing formula is as follows

$\mathrm{\Δ\α}{\mathrm{LT}}_c=\sqrt{{\left(L_x{\mathrm{\α}}_{\mathrm{cx}}\right)}^2+{\left(L_y{\mathrm{\α}}_{\mathrm{cy}}\right)}^2}\×{\mathrm{\ΔT}}_1$

$\mathrm{\Δ\α}{\mathrm{LT}}_s=\sqrt{{\left(L_x{\mathrm{\α}}_{\mathrm{sx}}\right)}^2+{\left(L_y{\mathrm{\α}}_{\mathrm{sy}}\right)}^2}\×{\mathrm{\ΔT}}_1$

In formula, L_xFor solder joint maximum spacing on x direction or pad array span, unit is mm；L_yFor solder joint on y direction Big spacing or pad array span, unit is mm；α_cxFor the thermal coefficient of expansion of device on x direction, unit is ppm/ DEG C；α_cyFor y The thermal coefficient of expansion of device on direction, unit is ppm/ DEG C；α_sxFor the thermal coefficient of expansion of circuit board on x direction, unit is ppm/ ℃；α_syFor the thermal coefficient of expansion of circuit board on y direction, unit is ppm/ DEG C.

Utilize Coffin-Manson heat exhaustion model, calculate Spacecraft Electronic assembly initial failure device thermal fatigue life N_f1, computing formula is as follows

$N_{f1}=\frac{1}{2}{\left(\frac{{\mathrm{\Δ\γ}}_1}{{2\mathrm{\ϵ}}_f^{\′}}\right)}^{\frac{1}{c}}$

In formula, N_f1For initial failure device thermal fatigue life under the conditions of thermal cycling test；Δγ₁For thermal cycling test condition Lower initial failure device solder joint shear strain；ε'_fFor fatigue ductile coefficient, generally take 2 ε '_f≈0.65；C is fatigue ductility index, meter Calculation formula is as follows

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

In formula, T_mFor test period solder joint mean temperature, unit is DEG C；t_DImmersing the time for maximum temperature, unit is min.

Judge that can Spacecraft Electronic assembly initial failure device screen out in thermal cycling test, i.e. thermal cycling test is alternative Whether the circulating cycle issue N of scheme is more than Spacecraft Electronic assembly initial failure device thermal fatigue life N_f1, statistics spacecraft electricity Sub-component can screen out initial failure device count m and initial failure device populations M, calculate Spacecraft Electronic component heat cyclic test Effectiveness E, computing formula is as follows

$E=\frac{m}{M}\×100\%$

If thermal cycling test effectiveness is 100%, then carry out step 4, if thermal cycling test effectiveness is not 100%, then Add tight condition, suitably increase maximum temperature T₁Or reduce minimum temperature T₂Or increase temperature cycles periodicity N, re-start step Rapid 2.

Wherein, Spacecraft Electronic assembly normal working temperature section described in step 4, Spacecraft Electronic assembly set Meter requirement determines, mainly includes the maximum temperature T in normally working₃, minimum temperature T₄, maximum temperature immerse time t₃, lowest temperature Degree immerses time t₄, rate temperature change v, the cycle projected life experience temperature cycles number N₁Deng.

The Spacecraft Electronic assembly digital prototype utilizing business software PWA, Flotherm etc. and step 2 to set up, counts respectively Calculate maximum temperature T during normal use₃And minimum temperature T₄In the case of, the highest temperature of Spacecraft Electronic assembly each device shell temperature Degree T '₃And minimum temperature T '₄, spacecraft electronic building brick each device temperature difference Δ T in normally being worked₂, the following institute of computing formula Show

ΔT₂=T '₃-T′₄

Wherein, described in steps of 5 Spacecraft Electronic assembly proper device solder joint effective area S₂Can be by normal solder joint Effective area actual measurement statistics or device handbook obtain.

Spacecraft Electronic assembly proper device solder joint shear strain Δ γ under the conditions of thermal cycling test₂, computing formula is as follows

${\mathrm{\Δ\γ}}_2=0.5\×\frac{K_D}{\left(1.38\mathrm{MPa}\right)S_{2}h}\×{({\mathrm{\Δ\αLT}}_s-{\mathrm{\Δ\αLT}}_c)}^2$

In formula, Δ γ₂For proper device solder joint shear strain under the conditions of thermal cycling test；K_DFor solder joint bending rigidity, unit is N/mm；S₂For proper device solder joint effective area, unit is mm²；H is solder joint height, and unit is mm；ΔαLT_cExpand for device Amount, Δ α LT_sFor circuit board swell increment, computing formula is as follows

$\mathrm{\Δ\α}{\mathrm{LT}}_c=\sqrt{{\left(L_x{\mathrm{\α}}_{\mathrm{cx}}\right)}^2+{\left(L_y{\mathrm{\α}}_{\mathrm{cy}}\right)}^2}\×{\mathrm{\ΔT}}_1$

$\mathrm{\Δ\α}{\mathrm{LT}}_s=\sqrt{{\left(L_x{\mathrm{\α}}_{\mathrm{sx}}\right)}^2+{\left(L_y{\mathrm{\α}}_{\mathrm{sy}}\right)}^2}\×{\mathrm{\ΔT}}_1$

In formula, L_xFor solder joint maximum spacing on x direction or pad array span, unit is in；L_yFor solder joint on y direction Big spacing or pad array span, unit is in；α_cxFor the thermal coefficient of expansion of device on x direction, unit is ppm/ DEG C；α_cyFor y The thermal coefficient of expansion of device on direction, unit is ppm/ DEG C；α_sxFor the thermal coefficient of expansion of circuit board on x direction, unit is ppm/ ℃；α_syFor the thermal coefficient of expansion of circuit board on y direction, unit is ppm/ DEG C.

Utilize Coffin-Manson heat exhaustion model, calculate the normal device of Spacecraft Electronic assembly under the conditions of thermal cycling test Part thermal fatigue life N_f2, computing formula is as follows

$N_{f2}=\frac{1}{2}{\left(\frac{{\mathrm{\Δ\γ}}_2}{{2\mathrm{\ϵ}}_f^{\′}}\right)}^{\frac{1}{c}}$

In formula, N_f2For proper device thermal fatigue life under the conditions of thermal cycling test；Δγ₂For under the conditions of thermal cycling test just Often device solder joint shear strain；ε'_fFor fatigue ductile coefficient, generally take 2 ε '_f≈0.65；C is fatigue ductility index, and computing formula is such as Under

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

In formula, T_mFor test period solder joint mean temperature, unit is DEG C；t_DImmersing the time for maximum temperature, unit is min.

Calculate the thermal cycling test test damage ratio D to Spacecraft Electronic assembly proper device₁, computing formula is as follows

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

Wherein, under described in step 6 spacecraft normal running conditions, Spacecraft Electronic assembly proper device solder joint is cut Strain Δ γ₃, computing formula is as follows

${\mathrm{\Δ\γ}}_3=0.5\×\frac{K_D}{\left(1.38\mathrm{MPa}\right)S_{2}h}\×{({\mathrm{\Δ\αLT}}_s-{\mathrm{\Δ\αLT}}_c)}^2$

In formula, Δ γ₃For proper device solder joint shear strain under normal running conditions；K_DFor solder joint bending rigidity, unit is N/ mm；S₂For proper device solder joint effective area, unit is mm²；H is solder joint height, and unit is mm；ΔαLT_cFor device swell increment, ΔαLT_sFor circuit board swell increment, computing formula is as follows

$\mathrm{\Δ\α}{\mathrm{LT}}_c=\sqrt{{\left(L_x{\mathrm{\α}}_{\mathrm{cx}}\right)}^2+{\left(L_y{\mathrm{\α}}_{\mathrm{cy}}\right)}^2}\×{\mathrm{\ΔT}}_2$

$\mathrm{\Δ\α}{\mathrm{LT}}_s=\sqrt{{\left(L_x{\mathrm{\α}}_{\mathrm{sx}}\right)}^2+{\left(L_y{\mathrm{\α}}_{\mathrm{sy}}\right)}^2}\×{\mathrm{\ΔT}}_2$

In formula, L_xFor solder joint maximum spacing on x direction or pad array span, unit is in；L_yFor solder joint on y direction Big spacing or pad array span, unit is in；α_cxFor the thermal coefficient of expansion of device on x direction, unit is ppm/ DEG C；α_cyFor y The thermal coefficient of expansion of device on direction, unit is ppm/ DEG C；α_sxFor the thermal coefficient of expansion of circuit board on x direction, unit is ppm/ ℃；α_syFor the thermal coefficient of expansion of circuit board on y direction, unit is ppm/ DEG C.

Utilize Coffin-Manson heat exhaustion model, calculate Spacecraft Electronic assembly proper device under normal running conditions Thermal fatigue life N_f3, computing formula is as follows

$N_{f3}=\frac{1}{2}{\left(\frac{{\mathrm{\Δ\γ}}_3}{{2\mathrm{\ϵ}}_f^{\′}}\right)}^{\frac{1}{c}}$

In formula, N_f3For proper device thermal fatigue life under normal running conditions；Δγ₃For device normal under normal running conditions Part solder joint shear strain；ε'_fFor fatigue ductile coefficient, generally take 2 ε '_f≈0.65；C is fatigue ductility index, and computing formula is as follows

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

In formula, T_mFor solder joint mean temperature under normal running conditions, unit is DEG C；t_DFor the highest temperature under normal running conditions Degree immerses the time, and unit is min.

Calculate the test damage ratio D to Spacecraft Electronic assembly proper device that normally works₂, computing formula is as follows

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

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

D=D₁+D₂

Judge whether Spacecraft Electronic assembly exists the life cycle management damage ratio D > 1 of proper device, if not existing, then Carrying out step 8, if existing, then softening terms, suitably reduce maximum temperature T₁Or increase minimum temperature T₂Or reduction temperature cycles Periodicity N, re-starts step 2.

Wherein, the thermal cycling test scheme the most finally determined is for meeting thermal cycling test effectiveness and thermal cycle The scheme that test acceptability requires, final output result is the maximum temperature T of thermal cycling test scheme₁, minimum temperature T₂, High-temperature immerses time t₁, minimum temperature immerse time t₂, 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 has: this Invent for concrete Spacecraft Electronic assembly, analyze the thermal cycling test thermal cycling test effectiveness to Spacecraft Electronic assembly And thermal cycling test is acceptable, more specific aim and reasonability.The present invention is theoretical based on Coffin-Manson thermal fatigue life With Miner linear cumulative damage law, draw effective and rational thermal cycling test scheme by computational analysis, have operable Property strong, save the advantages such as experimentation cost, can facilitate aerospace electron product thermal cycling test scheme of promptly formulating.

(4) accompanying drawing explanation:

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

In figure, label and symbol description are as follows:

T₁Maximum temperature for thermal cycling test

T₂Minimum temperature for thermal cycling test

t₁Maximum temperature for thermal cycling test immerses the time

t₂Minimum temperature for thermal cycling test immerses the time

N is the temperature cycles periodicity of thermal cycling test

V is the rate temperature change of thermal cycling test

T’₁For the maximum temperature of Spacecraft Electronic assembly each device shell temperature under the conditions of thermal cycling test

T’₂For the minimum temperature of Spacecraft Electronic assembly each device shell temperature under the conditions of thermal cycling test

ΔT₁Poor for each device temperature of Spacecraft Electronic assembly under the conditions of thermal cycling test

S₁For Spacecraft Electronic assembly initial failure device solder joint effective area

Δγ₁For Spacecraft Electronic assembly initial failure device solder joint shear strain under the conditions of thermal cycling test

K_DFor solder joint bending rigidity

H is solder joint height

ΔαLT_cFor device swell increment

ΔαLT_sFor circuit board swell increment

L_xFor solder joint maximum spacing on x direction or pad array span

L_yFor solder joint maximum spacing on y direction or pad array span

α_cxFor the thermal coefficient of expansion of device on x direction

α_cyFor the thermal coefficient of expansion of device on y direction

α_sxFor the thermal coefficient of expansion of circuit board on x direction

α_syFor the thermal coefficient of expansion of circuit board on y direction

N_f1For Spacecraft Electronic assembly initial failure device thermal fatigue life

Δγ₁For initial failure device solder joint shear strain under the conditions of thermal cycling test

ε'_fFor fatigue ductile coefficient

T_mFor test period solder joint mean temperature

t_DThe time is immersed for maximum temperature

M is to screen out initial failure device count in Spacecraft Electronic assembly

M is initial failure device populations in Spacecraft Electronic assembly

E is Spacecraft Electronic component heat cyclic test effectiveness

T₃For the maximum temperature in normal work

T₄For the minimum temperature in normal work

t₃The time is immersed for the maximum temperature in normal work

t₄The time is immersed for the minimum temperature in normal work

V is the rate temperature change in normal work

N₁Temperature cycles number is experienced for cycle projected life

T’₃For the maximum temperature of spacecraft electronic building brick each device shell temperature in normal work

T’₄For the minimum temperature of spacecraft electronic building brick each device shell temperature in normal work

ΔT₂Poor for each device temperature of spacecraft electronic building brick in normal work

S₂For Spacecraft Electronic assembly proper device solder joint effective area

Δγ₂For Spacecraft Electronic assembly proper device solder joint shear strain under the conditions of thermal cycling test

S₂For proper device solder joint effective area

N_f2For Spacecraft Electronic assembly proper device thermal fatigue life under the conditions of thermal cycling test

D₁Test damage ratio for Spacecraft Electronic assembly proper device

Δγ₃For spacecraft electronic building brick proper device solder joint shear strain in normal work

N_f3For spacecraft electronic building brick proper device thermal fatigue life in normal work

D₂For the work damage ratio of spacecraft electronic building brick proper device in normal work

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

(5) detailed description of the invention:

Below in conjunction with concrete case study on implementation, to the spacecraft for concrete Spacecraft Electronic assembly of the present invention Electronic building brick thermal cycling test scheme determines that method is described in detail.

Case: be applied to the rail control engine parameter collection plate of spacecraft

The present invention, as a example by the rail control engine parameter collection plate being applied to spacecraft, illustrates for concrete spacecraft The Spacecraft Electronic component heat cyclic test scheme of electronic building brick determines method.

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

(1) rail control engine parameter collection plate: this parameter acquisition plate is used for certain type geo-stationary orbit communications satellite, is positioned at and defends In the propelling module of star, a size of 96mm × 86mm × 1.53mm, total power consumption is 1.4W, comprises 11 electronic devices and components of 8 class, device Inventory is as shown in Table 4.

(2) application requires: this geo-stationary orbit communications satellite projected life is 15 years, and the temperature cycles cycle is 24 hours, its In, the temperature change time is 6 hours, and it is 6 hours that temperature immerses the time.Generally, the ambient temperature of propelling module is-10 to+40 DEG C, during collection plate work, the application of temperature environment requires as shown in Table 2.

Table two collection plate operating ambient temperature relevant parameter

According to the flow process of Fig. 1, the method specifically comprises the following steps that

Step 1: determine thermal cycling test alternative, primarily determines that the maximum temperature T of thermal cycling test₁It is 60 DEG C, Low temperature T₂For-25 DEG C, maximum temperature immerses time t₁For 60min, minimum temperature immerses time 60min, temperature cycles periodicity N is 25, and rate temperature change v is 4 DEG C/min.

Step 2: collecting Spacecraft Electronic device related information, the size of pcb board is 96mm × 86mm × 1.53mm, PCB The flaggy information of plate as shown in Table 3, the device item of each device in Spacecraft Electronic assembly, packing forms, material, position, The relevant information such as size, power consumption, 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
Third 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, calculate the maximum temperature T at thermal cycling test respectively₁And minimum temperature T₂The maximum temperature T ' of Spacecraft Electronic assembly each device shell temperature₁And minimum temperature T '₂, obtain space flight under the conditions of thermal cycling test Device electronic building brick each device temperature difference Δ T₁, as shown in Table 5.

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

Step 3: collect relevant parameter, calculate each device and circuit board swell increment thereof, as shown in Table 6.Utilize correlation formula And Spacecraft Electronic assembly initial failure relevant parameter, calculate Spacecraft Electronic assembly initial failure device under the conditions of thermal cycling test Part solder joint shear strain Δ γ₁, as shown in Table 7.Utilize Coffin-Manson heat exhaustion model, calculate aerospace electron assembly in early days Defective device thermal fatigue life N_f1, as shown in Table 8.

The each device of Spacecraft Electronic assembly and circuit board swell increment thereof under the conditions of table six thermal cycling test

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

Numbering	Item	ΔT1(℃)	ΔαLTc	ΔαLTs	KD	S1	h	Δγ1
									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	ΔT1(℃)	ΔαLTc	ΔαLTs	KD	S1	h	Δγ1
									11	D108	84.90089	0.0208	0.0587	2300	0.025893	0.25	0.2656

Table eight thermal cycling test efficiency analysis

Numbering	Item	ΔT1(℃)	Tm	tD	c	2ε'f	Δγ1	Nf1	＜ N
										1	C101	84.9656	17.8157	60	-0.418831	0.65	0.1965	8.7	It is
2	C102	84.9493	17.96507	60	-0.41892	0.65	0.1295	23.5	It is
										3	C103	84.93637	18.0966	60	-0.418999	0.65	0.1295	23.5	It is
4	D101	84.89719	18.74909	60	-0.419391	0.65	0.2684	4.1	It is
										5	D102	84.87448	18.98552	60	-0.419532	0.65	0.2772	3.8	It is
6	D103	84.95754	18.17347	60	-0.419045	0.65	0.1522	16.0	It is
										7	D104	84.96317	18.11839	60	-0.419012	0.65	0.1491	16.8	It is
8	D105	84.96311	18.13628	60	-0.419023	0.65	0.1603	14.1	It is
										9	D106	84.9197	18.79857	60	-0.41942	0.65	0.1875	9.7	It is
10	D107	84.94361	18.63799	60	-0.419324	0.65	0.1807	10.6	It is
										11	D108	84.90089	18.73037	60	-0.419379	0.65	0.2656	4.2	It is

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

Owing to Spacecraft Electronic component heat cyclic test effectiveness E is 100%, then carry out step 4.

Step 4: Spacecraft Electronic assembly normal working temperature section, is determined by the design requirement of Spacecraft Electronic assembly, main Including the maximum temperature T in normally working₃It is 50 DEG C, minimum temperature T₄For-10 DEG C, maximum temperature immerses time t₃It is 6 little Time, minimum temperature immerses time t₄Being 6 hours, rate temperature change v is 10 DEG C/h, and cycle projected life experience temperature is followed Number of rings N₁It it is 5475 times.

Utilize the business softwares such as PWA, Flotherm, calculate maximum temperature T in normal work respectively₃And minimum temperature T₄ The maximum temperature T ' of Spacecraft Electronic assembly each device shell temperature₃And minimum temperature T '₄, Spacecraft Electronic group in normally being worked Part each device temperature difference Δ T₂, as shown in Table 9.

During table nine normally works, each device temperature of spacecraft electronic building brick is poor

Step 5: collect relevant parameter, calculate each device and circuit board swell increment thereof, as shown in Table 6.Utilize correlation formula And spacecraft normal electrical assembly relevant parameter, under the conditions of calculating thermal cycling test, Spacecraft Electronic assembly proper device solder joint is cut Strain Δ γ₂, as shown in Table 10.Utilize Coffin-Manson heat exhaustion model, calculate aerospace electron under the conditions of thermal cycling test Assembly proper device thermal fatigue life N_f2, and calculate the thermal cycling test test damage ratio to Spacecraft Electronic assembly proper device D₁, as shown in Table 8.

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

Numbering	Item	ΔT1(℃)	ΔαLTc	ΔαLTs	KD	S2	h	Δγ2
									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

Table 11 thermal cycling test test damage ratio to Spacecraft Electronic assembly proper device

Numbering	Item	ΔT1(℃)	Tm	tD	c	2ε'f	Δγ2	Nf2	D1
										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 Spacecraft Electronic group under Spacecraft Electronic assembly normal solder joint relevant parameter and normal running conditions Part each components and parts temperature difference Δ T₂, calculate Spacecraft Electronic assembled devices thermal expansion amount and proper device weldering under normal running conditions Point shear strain Δ γ₃, as shown in table 12 and table 13.Utilize Coffin-Manson heat exhaustion model, calculate the bar that normally works Spacecraft Electronic assembly proper device thermal fatigue life N under part_f3, and it is normal to calculate Spacecraft Electronic assembly under normal running conditions The work damage ratio D of device₂, as shown in table 14.

Spacecraft Electronic assembled devices thermal expansion amount under table 12 normal running conditions

Spacecraft Electronic assembly proper device solder joint shear strain under table 13 normal running conditions

Numbering	Item	ΔT1(℃)	ΔαLTc	ΔαLTs	KD	S2	h	Δγ3
									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 14 normal running conditions

Numbering	Item	ΔT1(℃)	Tm	tD	c	2ε'f	Δγ3	Nf3	D2
										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: Spacecraft Electronic component heat cyclic test acceptability is analyzed: the test damage ratio obtained according to step 5 D₁The work damage ratio D obtained with step 6₂, calculate the life cycle management damage ratio D of Spacecraft Electronic assembly proper device, sentence In disconnected spacecraft electronic building brick, whether the life cycle management damage ratio D of proper device > 1, as shown in table 15.Due to spacecraft electricity Sub-component does not exist the life cycle management damage ratio D > 1 of proper device, carries out step 8.

The life cycle management damage ratio D of table 15 Spacecraft Electronic assembly proper device

Numbering	Item	D1	D2	D	Whether ＞ 1
						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: this thermal cycling test scheme finally determined is and rational thermal cycle examination effective for parameter acquisition plate Proved recipe case, as shown in table 16.

Table 16 thermal cycling test scheme

Claims (9)

1. a Spacecraft Electronic component heat cyclic test scheme determines method, it is characterised in that: utilize heat analysis, Coffin- Manson model and life cycle theory, acceptable based on thermal cycling test effectiveness and thermal cycling test, analyze heat Cyclic test scheme is to specifying the effectiveness of Spacecraft Electronic assembly and damage, and according to this method flow process, iteration judges, the most really Fixed suitable Spacecraft Electronic component heat cyclic test scheme, the method specifically comprises the following steps that

Step 1: the determination of thermal cycling test alternative；

Step 2: Spacecraft Electronic assembly each components and parts temperature difference Δ T under the conditions of thermal cycling test₁Calculate；

Step 3: Spacecraft Electronic component heat cyclic test efficiency analysis, if thermal cycling test effectiveness is 100%, is then carried out Step 4, if thermal cycling test effectiveness is not 100%, then adds tight condition, re-starts step 2；

Step 4: Spacecraft Electronic assembly each components and parts temperature difference Δ T under normal running conditions₂Calculate；

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

Step 6: Spacecraft Electronic assembly normally works breakdown diagnosis；

Step 7: whether Spacecraft Electronic component heat cyclic test acceptability is analyzed, it is judged that just exist in Spacecraft Electronic assembly Often the life cycle management damage ratio D > 1 of device, if not existing, then carries out step 8, if existing, then softens terms, re-starts step Rapid 2；

Step 8: the thermal cycling test scheme finally determined is the effective and rational thermal cycle for concrete Spacecraft Electronic assembly Testing program.

Spacecraft Electronic component heat cyclic test scheme the most according to claim 1 determines method, it is characterised in that: in step The main of thermal cycling test scheme described in rapid 2 determines that parameter is maximum temperature T₁, minimum temperature T₂, maximum temperature immerse the time t₁, minimum temperature immerse time t₂, temperature cycles periodicity N, rate temperature change v.

Spacecraft Electronic component heat cyclic test scheme the most according to claim 1 determines method, it is characterised in that: in step Utilize heat to analyze business software described in rapid 2, calculate each components and parts of aerospace electron assembly under each temperature levels of thermal cycling test Shell temperature, and draw shell temperature difference T of each components and parts of aerospace electron assembly₁。

Spacecraft Electronic component heat cyclic test scheme the most according to claim 1 determines method, it is characterised in that:

(1) the most described Spacecraft Electronic assembly initial failure device solder joint effective area S₁Can be real by defect solder joint area Statistics is measured on border or engineering experience obtains；

(2) under the conditions of the most described thermal cycling test, Spacecraft Electronic assembly initial failure is whole notes of device fault Enter；

(3) Spacecraft Electronic assembly initial failure device solder joint shear strain Δ γ under the conditions of the most described thermal cycling test₁ Computing formula

${\mathrm{\Δ\γ}}_1=0.5\×\frac{K_D}{\left(1.38MPa\right)S_{1}h}\×{({\mathrm{\Δ\αLT}}_s-{\mathrm{\Δ\αLT}}_c)}^2$

In formula, Δ γ₁For the solder joint shear strain of initial failure device, K under the conditions of thermal cycling test_DFor solder joint bending rigidity, unit is N/mm, S₁For initial failure device solder joint effective area, unit is mm², 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 mm；

(4) the most described Spacecraft Electronic component heat cyclic test effectiveness E computing formula

$E=\frac{m}{M}\×100\%$

In formula, E is Spacecraft Electronic component heat cyclic test effectiveness, and m is to screen out initial failure in Spacecraft Electronic assembly Device count, M is initial failure device populations in Spacecraft Electronic assembly；

(5) if described thermal cycling test effectiveness is 100% in step 3, then step 4 is carried out, if thermal cycling test effectiveness It is not 100%, then adds tight condition, suitably increase maximum temperature T₁Or reduce minimum temperature T₂Or increase temperature cycles periodicity N, re-starts step 2.

Spacecraft Electronic component heat cyclic test scheme the most according to claim 1 determines method, it is characterised in that: in step Utilize heat to analyze business software described in rapid 4, calculate each components and parts of aerospace electron assembly and work under each temperature levels normal Shell temperature, and draw shell temperature difference T of each components and parts of aerospace electron assembly₂。

Spacecraft Electronic component heat cyclic test scheme the most according to claim 1 determines method, it is characterised in that:

(1) the most described Spacecraft Electronic assembly proper device solder joint effective area S₂Can be real by normal solder joint effective area Statistics is measured on border or device handbook obtains；

(2) Spacecraft Electronic assembly proper device solder joint shear strain Δ γ under the conditions of the most described thermal cycling test₂Calculate Formula

${\mathrm{\Δ\γ}}_2=0.5\×\frac{K_D}{\left(1.38MPa\right)S_{2}h}\×{({\mathrm{\Δ\αLT}}_s-{\mathrm{\Δ\αLT}}_c)}^2$

In formula, Δ γ₂For the solder joint shear strain of initial failure device, K under the conditions of thermal cycling test_DFor solder joint bending rigidity, unit is N/mm, S₂For proper device solder joint effective area, unit is mm², h is solder joint height, and unit is mm, Δ α LT_cExpand for device Amount, unit is mm, Δ α LT_sFor circuit board swell increment, unit is mm；

(3) the most described thermal cycling test test damage ratio D to Spacecraft Electronic assembly proper device₁Computing formula

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

In formula, D₁For the thermal cycling test test damage ratio to Spacecraft Electronic assembly proper device, N is the temperature of thermal cycling test Spend circulating cycle issue, N_f2For proper device thermal fatigue life under the conditions of thermal cycling test.

Spacecraft Electronic component heat cyclic test scheme the most according to claim 1 determines method, it is characterised in that:

(1) Spacecraft Electronic assembly proper device solder joint shear strain Δ γ under the most described spacecraft normal running conditions₃ Computing formula

${\mathrm{\Δ\γ}}_3=0.5\×\frac{K_D}{\left(1.38MPa\right)S_{2}h}\×{({\mathrm{\Δ\αLT}}_s-{\mathrm{\Δ\αLT}}_c)}^2$

In formula, Δ γ₃For the solder joint shear strain of initial failure device, K under the conditions of thermal cycling test_DFor solder joint bending rigidity, unit is N/mm, S₂For proper device solder joint effective area, unit is mm², h is solder joint height, and unit is mm, Δ α LT_cExpand for device Amount, unit is mm, Δ α LT_sFor circuit board swell increment, unit is mm；

(2) the most described normal work test damage ratio D to Spacecraft Electronic assembly proper device₂Computing formula

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

In formula, D₂For the test damage ratio to Spacecraft Electronic assembly proper device that normally works, N₁Experience for cycle projected life Temperature cycles number, N_f3For proper device thermal fatigue life under normal running conditions.

Spacecraft Electronic component heat cyclic test scheme the most according to claim 1 determines method, it is characterised in that:

(1) the life cycle management damage ratio D computational methods of the most described Spacecraft Electronic assembly proper device

D=D₁+D₂

In formula, D is the life cycle management damage ratio of Spacecraft Electronic assembly proper device, D₁For thermal cycling test to spacecraft electricity The test damage ratio of sub-component proper device, D₂For the test damage ratio to Spacecraft Electronic assembly proper device that normally works；

(2) the most described life cycle management damage ratio D judging whether to there is proper device in Spacecraft Electronic assembly > 1, if not existing, then carrying out step 8, if existing, then softening terms, suitably reduce maximum temperature T₁Or increase minimum temperature T₂ Or reduce temperature cycles periodicity N, re-start step 2.

Spacecraft Electronic component heat cyclic test scheme the most according to claim 1 determines method, it is characterised in that: in step Spacecraft Electronic component heat cyclic test scheme described in rapid 8 is the maximum temperature T of thermal cycling test scheme₁, minimum temperature T₂、 Maximum temperature immerses time t₁, minimum temperature immerse time t₂, temperature cycles periodicity N, rate temperature change v.