Design-for-Reliability Through Durability Analysis

Wilbur W. Bhagat; Aeronautical Systems Division, WPAFB

Bruce A. Tagg; Aeronautical Systems Division, WPAFB

Key Words: Design for Reliability, Durability, Design Criteria,

Materials/Parts Characterization, Service Life,

Design Life, Fatigue, Cumulative Damage, Avionics

Summary and Conclusions

This paper discusses an analytical technique called "Durability Analysisv1 which, when applied to avionics during the design process, will improve or enhance the avionic hardware reliability and durability. This improvement or enhancement in reliability/ durability occurs early in the development phase by way of design changes on paper and by consideration of various design trade-offs based on criticality of the hardware and on economics of maintenance support. This paper discusses various types of durability analysis with some examples. It also includes guidance on which areas or parts of the hardware should be selected for analysis.

The durability analyses performed for the current avionics acquisition programs managed by the Aeronautical Systems Division (ASD) identified the potential reliability problems early in the design phase. They were corrected by changing the design on paper. This avoided or reduced the potential for costly redesign and retrofit efforts and the ssociated maintenance burden which would have occurred had these corrective actions been implemented after the release of the design for fabrication. This paper presents the types of design changes and corrective actions that were identified by the durability analysis.

1. Introduction

Modern avionics contain a large number of highly sophisticated and complex electronic hardware. Demand for more complex functions in conjunction with the need for lighter and more compact packaging has increased the potential for reliability problems. Therefore the need for achieving reliability in a timely and cost effective manner has become increasingly important for avionic hardware. As discussed in Ref. 1, the traditional reliability prediction, analysis and test techniques had limited success in achieving reliability in a cost effective manner early in the design and development phase. To address this problem, ASD in the mid-eighties initiated a technical approach called "Avionics/Electronics Integrity Program (AV1P)Il for the new avionics acquisition programs. An overview of this approach was presented at the 1989 AR&MS (Ref. 1). AVIP has been derived from, and its structure is based upon, other successful integrity

programs, i.e., Aircraft Structural Integrity Program (ASIP) and Engine Structural Integrity Program (ENSIP). The key elements of AVIP include understanding and establishing the design usage and environments which cause stresses on the avionic hardware, understanding and establishing the design usage and environments which cause stresses on the avionic hardware, understanding and determining the properties and characteristics of the parts, materials and processes that will be used in the avionics, durability analysis, selecting and performing engineering/development tests, qualifying the hardware for the required service life by performing a durability life test (DLT) , and establishing provisions for quality assurance and life management: of these elements, i.e., durability.

The term "Durability analysis" refers to a set or group of analyses that are performed to identify durability/reliability problems in the avionic hardware and to predict life of the hardware for given design usage and environments. These analyses are performed using analytical techniques, formulas and equations from various technical disciplines such as mechanical engineering, electrical engineering, materials engineering, strength of materials, structural engineering, etc.

This paper addresses one

The next section explains the terms which are unique to integrity programs or which may require explanation to facilitate better understanding of the material presented in this paper. The subsequent sections will discuss tasks which are prerequisite to durability analysis, different types of durability analysis, the analytical tools presently available and the results from the analyses performed recently. The appendix shows examples of the analysis.

2. Nomenclature

a. Durabilfty: The ability to resist cracking, corrosion, deterioration, Thermal degradation, delamination, wear, electro- migration, dielectric breakdown, and other failure modes for a specified period.

Hardware material element (e.g., lead wire, solder joint, component, etc.) the life characteristics of which control and determine the durability and life of the hardware.

b. Durability Control Point (DCP):

516 0149-144X/91/0000-0516$01.00 0 1991 IEEE

1991 PROCEEDINGS Annual RELIABILITY AND MAINTAINABILITY Symposium

3 . Prereauisties for Durabilitv Analvsis The main objective of the durability

design process is to prevent failures from occurring during the operational service life of the equipment, or during an agreed upon period of operational usage. this, it is necessary to understand major expected or potential failure modes that may occur during this time period. These failures are usually caused by stresses which are imposed on various parts or areas of the hardware due to given usage and environments. The stress types, levels and durations will depend upon how the avionic hardware will be used and the environments in which it will be used. Therefore, it is very important to understand and establish the design usage and environmental profiles for the equipment early in the design and development phase. information will be necessary for performing the durability analyses.

Another important task which should be initiated prior to the durability analysis effort is materials and parts characteri- zation. This task entails establishing, obtaining or determining various properties and characteristics of the materials and parts that will be used in the avionic equipment. This information will facilitate durability and other engineering analyses, and will help the designer determine the capability and suitability of various materials and parts for a given application. and part characteristics can be obtained from literature. Examples of materials/part characterization can be found in Ref. 2-6. When the data is not available, small-scale testing (sample or coupon) in many cases can provide the necessary data.

4 . gurabilitv Analvsis Process

To accomplish

This

Many of these material

The purpose of the durability analysis is to ensure that the avionic equipment will perform as specified for the required service life when subjected to the specified service usage and environments which include flight operation, ground operation, and logistic environments. The durability analysis is also used to verify the design and derating criteria established earlier in the design phase. The analysis usually addresses major stresses which are expected to occur in the equipment during its life cycle and which, if not controlled, cause failures in the equip- ment. There are two types of stresses that the avionic equipment usually experiences: steady stress (e.g., time at temperature of at power) and cyclic stress (e.g., vibration, thermal cycling).

are established, the duty cycles and enviromental profiles for the avionics are de- rived. This data then can be used to determine the stresses that the avionics will experience during various phases of the life cycle. A Failure Modes, Effects and Criticality Analysis FMECA), using MIL-TSD- 1629 as a guide, will identify most of the major failure modes. Knowing the stresses and the possible failure modes, one can use appropriate life laws for those failure modes or mechanisms, and estimate expected life of

Once the design usage and environments

the equipment for the given usage and environment.

The failure process in avionics can generally be categorized into three major categories: electrical, mechanical, and chemi- cal. The durability analysis process will be discussed below in terms of these three major categories.

a. Electrical: Electrical failures are usually initiated or caused by electrical stresses such as current or voltage. However, the actual physical mechanism of failure may be mechanical or structural in nature. Examples are electromigration, time dependent dielectric breakdown (TDDB), etc. Life laws for many of these failure modes have been published in the literature such as Reliability Physics Symposium. The following is an example of a life equation for electromigration in an electrical conductor. (Ref 7).

where t = m time to failure

A = cross-sectional area of the Conductor, an2

J = current density, -/an2

E = Activation energy, ev

K = Boltman's Constant 8 . 6 2 ~ 1 0 - ~

T = mrature OK

eV/OK

Two examples of TDDB are shown in Ref. 8 (Intel model) and Ref. 9 (Berkeley model). The Berkeley model for Time to Breakdown is as follows: ..

where v = gate voltage stress, V BD = Breakdo% voltage of a ramp-voltage stress with ramping rate R, B and H are the exponen- tial factors for the Fowler-Nordheim current.

For the failure modes for which the reliability physics laws or models cannot be found in the literature, and for which no experimental, test or field experience data exists, small-scale testing can be performed on coupons, samples, or mockups in order to predict the durability of the hardware for the subject failure mode. Some of these tests may be accelerated.

b. Mechanical: Thermal cycling and vibration are two major types of mechanical stresses in modern avionics. These stresses are cyclic in nature and cause fatigue failures in the material elements of avionics such as solder joints, component leads, plated-through-holes, etc. The fatigue caused by thermal cycling is called low cycle fatigue, and the fatigue caused by vibration is called high cycle fatigue.

1991 PROCEEDINGS Annual RELIABILITY AND MAINTAINABILITY Symposium 517

(1) Thermal (or low cycle) fatigue: Thermal cycling at a given point in avionics can occur due to either environmental temperature cycling or the device power cycling, or both. In all of these cases, a transient thermal analysis is conducted to determine the thermal cycling response at the durability control point (DCP). The number of thermal cycles (nl, n2,..) and the temperature excursions (ATl, AT 2 1 . . . ) expected during various phases of the equipment life cycle are estimated. From the fatigue characteristic of the DCP material (Ref. 1-6, lo), the number of cycles to failure (N1, N2,...) are calculated. The ratios %,” damage ratio& 8% fatigue damage indices. sum of all these ratios (Da )represents the cumlative damage developed Suring thermal cycling. It is also representative of the fraction of the useful life used up by thermal cycling .

fatigue: The majority of modern avionic equipment experience random vibration. The vibration response spectra at the DCP are obtained by performing vibration analysis or test measurements.

, etc., are called fatigue The

(2) Vibration (or high cycle)

The number of fatigue cycles accumulated during various phases of the equipment life cycles are calculated using the natural frequency of the printed circuit assembly (PCA) and assuming a Gaussian distribution for vibration energy levels. Ref. 10 shodthis calculation technique. From the high cycle fatigue characteristic for the DCP material the number of fatigue cycles to failure are obtained for each vibration condition. As was done for thermal fatigue, the fatigue damage ratios are computed for each vibration condition. The sum of these ratios represents cumulative damage (hIB) developed during vibration.

The sum D = D A T represents the total fatigue damage tt?% equipment will experience during one service life. The contributors to this fatigue damage include thermal cycling and vibration during Environmental Stress Screening (ESS) and during flight operations, thermal cycling during performance checks on ground (on air vehicle and off the air vehicle), storage diurnal cycles, vibration during transportation, thermal cycling due to solder- desolder repair events, etc. The modified Miner’s rule, derived from the air frame and engine experience and studies, indicates that there is less likelihood of fatigue failure if Dh0.7. For the durability design, a design margin of 2 is recommended. The design criterion then becomes DS0.35.

It should be noted that this cumulative damage technique is a linear first order approximation. It is a very helpful tool in comparing various fatigue stress conditions for their degree of severity. When D is very high, it raises a flag to indicate that there are some potential problem areas in the hardware design that may require additional or more sophisticated analysis and/or testing.

c. Chemical: The chemical failure

process in avionics mainly refers to the corrosion process. Corrosion is a gradual process of chemical erosion of materials, which is accelerated by the presence of moisture, contamination, and heat. During the design process, there are various considerations and criteria undertaken and implemented to prevent or control corrosion in avionics. There are some quantitative (analytical) tools available in literature for certain types and conditions of corrosion. Many of the corrosion prevention and control criteria are based on previous experience and experimental data. Some major considerations include:

- - contamination control with respect to

proper selection of materials and dimensions

soldering flux residue, voids in conformal coating, etc.

protective coating such as paint

minimizing exposure to moisture and other corrosive environments

- -

5. Other Considerations for Analvsis

important to identify and consider the effects of variability associated with manufacturing and assembly, on the equipment life. Many stresses that limit the equipment life, are influenced by this variability. of the variability are as follows:

a. Material variability (e.g., fatigue characteristics, thermal expansion coefficients, dispersion in electromigration or TDDB life, etc.)

off height of interconnections, thickness of dielectric barriers, lead dimensions etc.)

For the safety-critical and mission critical items, the worst case conditions associated with all applicable variabilities are addressed in the analysis.

During the analysis process, it is

Some examples

b. Dimensional variability (e.g., stand-

6. Analvtical Tools A large part of durability analysis

consists of “order of Magnitude” type calculations. Many of these are first order approximations. analysis include Strength of Materials equations (e.g., stress, bending moment, deflection, etc.), basic engineering equations from various engineering disciplines, laws of physics, reliability physics models, cumulative fatigue damage technique, modified Goodman diagrams, etc.

Personal Computers (PC) provide the next level of sophistication. Spreadsheet software can be used to accelerate the analysis, and also can be used as a powerful tool to evaluate various design options, especially when considering variabilities of the durability critical parameters.

the next level of sophistication. A number of them with different capabilities are now

The tools for this type of

Finite element analysis programs provide

1991 PROCEEDINGS Annual RELIABILITY AND MAINTAINABILITY Symposium

-

available for both PCs and mainframe computers. However, care must be exercised when using them because of the micro-structure of the avionics.

7. Current Exlserience

Durability analyses were performed on a number of acquisition programs managed by ASD, as one of the AVIP requirements. These analyses identified a number of design changes or modifications that were implemented on paper design before the design release for fabrication. Had these changes or modifications been identified and accomplished later on during qualification testing or in the field, they would have adversely impacted the program cost and schedule, and the available design options would have been much more limited. The following are examples of the types of design modifications that resulted from the durability analyses.

Component location changes

Component lead configuration modification

Component mounting modification

Board location changed

Hybrid stand-off height adjusted

Board configuration/composition changed

Heat sinks changed

Board installation modified

Process controls established

The fatiaue characteristics of many materials use; in avionics are expressei in terms of stress or strain versus number of cycles to failure curves. These are called S- N curves. In order to determine the number of cycles to failure for a given condition of usage and environment, one needs to know the stress (or strain) associated with that condition. This stress can be calculated knowing the material properties and geometry, and using the basic strength of materials and engineering equations. Two example calculations are included in Figures 1 and 2.

Figure 1 shows a surface mount choke under thermal cycling environment. o< Values are coefficients of thermal expansion. A T represents temperature excursion. Other parameters such as E, I, and C, are the same as those used in standard strength of materials equations.

leaded component such as transformer or choke, with printed circuit board (PCB), under vibration. F is the natural frequency of the PCB. Z is maximum deflection of the PCB. K is &e spring constant of the lead wire.

Figure 2 shows poke-thru assembly of a

References

1. W. Bhagat, "R&M Through Avionics/ Electronics Integrity Programm1 , Proc. Ann. Reliability and Maintainability Symp., 1989 January, pp 216-220.

IIThermal Fatigue in Silicon Power Transistorsll, IEEE Transactions on Electronic Devices, 1970 September.

3. D. Steinberg, Wibration Analysis for Electronic Equipment,lI 1988, Second Edition; John Wiley & Sons.

4. Wright Research & Development Laboratories Technical Report, *Vibration Stress Analysis of Avionics", AFWAL-TR-87-3023, 1987 April.

2. G. Lang, B Fehder, W. Williams,

5. Wright Research & Development Laboratories Technical Report, IIVibration Reliability Life Model for Avionics", AFWAL-TR-87-3048, 1987 September.

6. R. Tummala, E. Rymaszewski, lpMicroelectronics Packaging Handbook1', 1989; Van Nostrand Reinhold.

7. J. Partridge, A Marques, R. Camp, llElectromigration, Thermal Analysis and Die Attach - A Case History1', Proc. Int. Reliability Physics Symp., 1982.

Wearout Due to Charge Trappingm1, Proc. Int. Reliability Physics Symp., 1983.

Quantitative Physical Model for Time-Dependent Breakdown in Si0 Proc. Int. Reliability Physics Symp., 18i5.

Implementing AVIPI', Proc. IEEE/AESS Dayton Chapter Conference, 1988 November.

8. W. Meyer, D. Crook, IlModel for Oxide

9. I. Chen, S. Holland, C. Hu, "A

10. D. Steinberg, llTools Available for

BIOGRAPHY

Wilbur W. Bhagat, PE

ASD/ENASV

Wright-Patterson AFB, OH 45433-6503 USA

Wilbur Bhagat is a senior Avionics Integrity Engineer and a Group Leader in the Avionics Integrity Branch of the Aeronautical Systems Division (Air Force Systems Command). In this capacity, he has addressed durability/ reliability and other product integrity issues for a number of Air Force systems including LANTIRN, ATARS, AN/ALE-47 CMDS, F-111 Digital Flight Control System, Tail Warning System, etc. He holds an MSME degree from the University of Arkansas, and is about to complete the graduate work for an MS in Engineering Management at the University of Dayton. He is a registered Professional Engineer.

1991 PROCEEDINGS Annual RELIABILITY AND MAINTAINABILITY Symposium 519

Bruce A. Tagg

AS D/ENASV

Wright-Patterson AFB, OH 45433-6503 USA

Integrity Branch at the Aeronautical Systems Division (Air Force Systems Command), Wright- Patterson AFB, Ohio, as an Avionics Integrity Engineer. In this capacity, he support F-15EI Mark XV, and Short Range Attack Missile programs. Lt Tagg attended the University of Michigan at Ann Arbor, Michigan, and received a BS degree in Aerospace Engineering in 1988.

Bruce Tagg is assigned to the Avionics

4 - - + h

4 I I < cl t L1

520

Epoxy PCB: Yi = ai Li AK

Iron Core Body: Y2 = crz L2 AT2

Copper Wire: YJ = cu L, ~ T J

Relative Displacement of Wire: 6 = YC - (YP + Y3)

Horizontal Force In Wire: 6 E I

H = I! [1/3 - h/(4h+L)]

Bending Moment in Wire: M = Hh[l -2h/(4h + L)] Wire Bending Stress: Sb Mc/l

Solder Shear Stress: st = H/A

Figure 1

r"'/"l c

Z - P . ' G Z 2,SIN 0- FR

Zi=Ze-Z p = k 6 k = y

TENSILE STRESS IN WIRE: S,= f SHEAR STRESS IN SOLDER: S, = s,

WHERE 4 E X L t c

Figure 2

1991 PROCEEDINGS Annual RELIABILITY AND MAINTAINABILITY Symposium