29 June 2010

An Efficient Theoretical Technique To Predict Fatigue Failure In Satellite

Structure Before Qualification Testing

A.M. Elhady1, Dalia N. Ahmed2, Somaia T. Ahmed2, and Karim A. Abd Elrazik2.

A. M. ElhadyHead of Testing and Quality Sector

Head of Mechanical Group Egyptian Space Program

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Outline

IntroductionEgyptsat-1 SatelliteSatellite Structure DescriptionPerformance Requirements and Mechanical LoadsModel Description and AnalysisConclusions

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Introduction

The present work describes the theoretical and experimental design analysis carried out on Egyptsat-1 satellite structure. The design analysis here supports the qualification process and aims to ensure the satellite structure can survive and resist the expected mission environments and perform its function safe through the overall satellite mission life.Theoretical analysis is conducted by means of the finite element method for static and dynamic cases. The calculated g RMS values for the static analysis are applied in each axis of the satellite structure assembly. The dynamic case is used to determine the modal shapes and resonance frequencies. The stress values are calculated for the applied static and dynamic load cases as well as the fatigue analysis.

Egyptsat-1 Satellite

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Figure 1. General exploded view of the satellite structure.

Structure Description

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Figure 2. Frame Module General View

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Figure 3. General exploded view the upper plate

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Figure 4. General exploded view of the bottom plate.

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Figure 5. General exploded view of the precise mounting unit.

Performance Requirements and Mechanical LoadsThe satellite structure, as an integral part of the spacecraft, shallbe capable for supporting the spacecraft total mass of 160 Kg and,in order to guarantee the controllability of the launcher.the satellite fundamental frequencies, considering the adapter inthe frequency determination, shall be:Minimum transverse: 10 Hz;Minimum longitudinal: 20 Hz;The structure should keep all its specified dimensions, alignmentsand tolerances, taking into account the required positioningaccuracy of the attitude determination and control sensors andactuators relative to the payload during the mission lifetime afterbeing subjecting to all handling, testing, transportation, launchingand orbital loads.

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Mechanical load in each phase defined according to statisticalmeasurements done for long time.Mechanical loads affecting the satellite are prescribed for variousphases of satellite realization. First phase is the transportation fromthe AI&T facility to the launch site by one of the following:

Road transportation;Rail transportation; orAir transportation.

Second phase is the launching and orbit injection. Mechanical loadsof this phase are predefined in launch vehicle user manual.

Model Description and Analysis

Static AnalysisIn the static analysis, the component mass has beenapplied approximately at their locations, and thesatellite launch vehicle interface three points realizedon the surface of upper plate have used as fixed pointsfor boundary conditions. After several runs, theoptimum values for the discussed geometricparameters have been finalized with a safety marginsno less than 0.3 to no more than 5.

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Theoretical Dynamic Analysis1) Using passive or active damping; In this method we could use materials with

high damping coefficient or in active case, actuators are used.2) Controlling and decreasing dynamic response of structure in sensitive

sections; in this method, actuators are used in specified sections, which arevery sensitive to vibration and displacement.

3) Frequency separation; this method is one of the most common approachesused for vibration control. In this method the designer tries to shift the naturalfrequency of the structure to a safe range far from the excitation loadsfrequency.

4) Controlling the applied load function; in this method we try to control theapplied load to not happen in the same frequency as the system first naturalfrequency appears.

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Based on the FE model for preliminary designstage that has been used for static analysis aseries of modal analyses have been conductedby using ANSYS FE code.Imagers and attitude determination and controlsensors and actuators mounted on precisemounting unit were simulated by concentratedmasses at their center of mass has equivalentmoments of inertia.

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As a result of these series of modalanalyses the satellite structure naturalfrequencies lower than 100 are 43, 50, 80,100.

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Experimental Static AnalysisSatellite Strength mock-up was installed ontechnological bracket that allows performing of theturning over of the strength mock-up from verticalposition at an angle in accordance withcorresponding test set-up as shown in figure.

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Margin of safety of satellite structure units was definedaccording to the formula:

11

222

2

−++

+=

zyx

t

nnn

nη

Margin of safety calculated at the center of componentmounted on the satellite body was minimum at thebottom plate with value 0.25, while it increasesgradually moving to the upper plate with safety marginof 0.59.

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Experimental Dynamic AnalysisExperimental dynamic testing of satellite strengthmock-up aims at verifying that the satellite flightmodel structure shall bear up all induced loads fromthe moment of manufacturing up to the end of itsservice life within the permissible deformation.Satellite strength mock-up vibration tests werecarried out by shuttle scanning with a speed of 0.5octaves per minute.

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During air transportation vibration testalong Y1-axis over frequency band 10-25Hz after test conducting by 9min.Visual inspections of the satellite strengthmock-up after air transportation testsshowed two cracks on the bracket forMBEI OMU and one crack on the plate ofthe precise mounting unit.

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Figure 7 – Photos of the crack in the plate of precise mounting unit

Figure 8 – Photos of the crack in bracket for attachment MBEI OMU

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Analysis of the recorded experimental acceleration values atshuttle scanning test showed that structural transmissibilitycoefficient is approximately 19 at 17 Hz as shown in figure.

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This means that the theoretical simulation of the satellitestructure was lacked for accuracy. The precise mountingunit re-simulated as shown in figure.

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Series of modal analyses have been conducted by usingANSYS FE code. Frequency swab from 9-100 Hz with2Hz step, are applied to the new simulated configurationand maximum calculated equivalent Von-Misses stress arenormalized at each frequency. Comparison of thecalculated normalized equivalent stresses to theexperimental transmissibility coefficient presented infigure.

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Conclusions

The static analysis results shows that a high value ofsafety factor has been considered, but it was proved thatfor small satellite, dynamic design consideration havethe main rule on the structure capability under appliedload.The tasks which are obtaining natural frequencies,recognition of frequency range, and the nature of theapplied dynamic loads have been explained. These tasksare necessary for avoiding resonance phenomenon inthe structure.

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Parametric design approach was used to model geometry to minimize the design loops for optimization.The frequency swap method using ANSYS FE code,shows reasonable agreement with the experimental results.More analysis is required to support the mathematical baseof the frequency swab method and approve it astheoretical technique for satellite structure analysis.

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Thank You

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