Nationaal Lucht- en Ruimtevaartlaboratorium National Aerospace Laborator y NLR

NLR-TP-98264

Failure criterion for the skin-stiffener interface in composite aircraft panelsJ.C.F.N. van Rijn

Nationaal Lucht- en Ruimtevaartlaboratorium National Aerospace Laboratory NLR

NLR-TP-98264

Failure criterion for the skin-stiffener interface in composite aircraft panelsJ.C.F.N. van Rijn

This investigation has been carried out under a contract awarded by the Netherlands Agency for Aerospace Programmes, contract number 01310N. The Netherlands Agency for Aerospace Programmes has granted NLR permission to publish this report. This report is based on an article to be published in the proceedings of the Thirteenth Technical Conference of the American Society for Composites.

Division: Issued: Classification of title:

Structures and Materials 10 June 1998 Unclassified

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ContentsABSTRACT INTRODUCTION EXPERIMENTAL INVESTIGATION SPECIMEN DESCRIPTION EXPERIMENTAL PROCEDURE EXPERIMENTAL RESULTS RESULTS FROM EARLIER INVESTIGATIONS Specimen description Experimental procedure Experimental results EVALUATION OF EXPERIMENTAL RESULTS OBSERVATIONS AND INTERPRETATION OF EXPERIMENTAL RESULTS CRITERION DEFINITION CONCLUSIONS EXPERIMENTAL RESULTS FROM THE LITERATURE CONCLUSIONS AND RECOMMENDATIONS CONCLUSIONS RECOMMENDATIONS ACKNOWLEDGEMENTS REFERENCES 5 5 6 6 6 7 8 8 8 8 8 8 9 11 11 13 13 13 14 14

6 Tables 17 Figures

(33 pages in total)

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TEST AND VERIFICATION EQUIPMENT FOR THE ATTITUDE & ORBIT CONTROL SYSTEM OF THE XMM SATELLITE Contents by H.A. van Ingen Schenau, L.C.J. van Rijn, J. Spaa FAILURE CRITERION FOR THE SKIN-STIFFENER Abstract INTERFACE IN COMPOSITE AIRCRAFT PANELS National Aerospace Laboratory NLR

3 3

1.ABSTRACT 2.

J.C.F.N. van Rijn E-mail: ingen@nlr.nl; lvrijn@nlr.nl; spaa@nlr.nl National Aerospace Laboratory NLR 3 Tensor Skin Concept nature of the Amsterdam, The NetherlandsAOCS. A complete AOCS, together with

Introduction Tel. +31 527 248 218; Fax: +31 527 248 210

P.O. Box 153, 8300 AD Emmeloord, The Netherlands

dynamics and environment can be considered as a The National Aerospace Laboratory NLR in the loop which is actively closed by the Attitude Control 4 3. Experimental Results Netherlands has developed a new generation of Test Computer (ACC). and Verification Equipment (TVE) for testing of Attitude and 4. Simulations Orbit Control Subsystems of spacecraft. 7 target attitude ABSTRACT Based on a prototype TVE developed for ESA, test measurements commands equipment has been developed for Matra Marconi CONTR L O Space for AOCS subsystem performed Tests testing and tests it level established Developments system was and Future that, in the post-buckling regime, failure of 9 In 5. Recently numerous panel of a composite stiffened panel is often induced by failure of a skin-stiffener interface. The present the XMM and INTEGRAL scientific satellites.

9 This paper 6. Conclusions describes the test concept and the NLR, which aims at test system with its failure criterion for stiffener pop-off. architecture of the XMM the derivation of a main The the specimens used in the present investigation consisted of DYNAMICS torques a tapered stiffener flange magnetic attitude features, strip incremental development and delivery, Info on: fields 9 References and experiences aobtained during development and usewere loaded in four-point bending. bonded on skin laminate. The specimens disturbances Earth Sun of It was established that there washas significant influence of the stiffness properties of the flange the system. The described work no also been ONMENT performed under the skin-stiffener interface strength. The Stars field of theENVIR ESA contract. Magnetic laminate on strength skin-stiffener interface was

paper describes recent and on-going research performed at the National Aerospace Laboratory

SENSORS

ACTU TORS A

Syst

1. INTRODUCTION 4 Tables Fig. 1 Generic attitude contr ol system Based on experiences with the production and use of 8 Figures various test systems for the ISO, SAX, SOHO and other satellites, the National Aerospace Laboratory In the integration and test phase the AOCS subsystem INTRODUCTION Physical Le vel NLR in the Netherlands has developed a new is gradually built up Attitude depending on the schedule of Data generation In generic Test panel tests it was established incoming the post-buckling regime, control modes of numerous and Verification Equipment Contr l Computer o that, in units. Verification of attitude failure of AOCS (TVE) with re-usable hardware and software for and real-time behaviour is done in the early period of a composite stiffened panel is often induced by integration using a combination interface. It was failure of a skin-stiffener of real and simulated MA b CS us testing of Attitude and Orbit Control Subsystems therefore deemed necessary to include a failure criterion for the skin-stiffener interface as a (AOCS) of spacecraft [Ref. 1]. units. Protocol Le vel

governed by the deformation of the skin only.

Sy B Int

constraint in the panel design optimization code PANOPT. PANOPT was developed at the Sensor ActuatorElectr nics o Electr nics o

Data

National to be usable from the early stage of the The TVE hadAerospace Laboratory NLR for the design of stiffened composite panels for primary a The test concept described in this paper is based on AOCS development up to the integration and post-buckling constraints (Ref. 1). of the AOCS static closed loop test facility (no real motion). The aircraft structures with buckling Sensor in the spacecraft environment at the NLR, aimed at the development of a is showncriterionActuatorthedynamics i.e. open loop tests with test configuration failure in figure forThe skin2. Head Earlier research User v a single unit up to closed loop tests with any and environment simulation is responsible for the Leel stiffener interface, was reported in references 2 and 3. Data combination of real and simulated AOCS units should computation Electrical of stimuliCSfor the CS sensor Electrical and the units Physical MA unit MA In reference 2 the results of an experimental programme on strip specimensfrom thesecondarily (Engineer with a actuator units. ing units) be supported. processing of uli Stimmonitor ulation Stim uli Simdata Monitoring Monitoring bonded blade-type stiffener was presented. The specimens were tested in four point bending, The stimulated sensors will deliver sensor Gr of A under lateral tensionbuilt the skin, and with a pull-off test. AntoevaluationSim the results was prototype TVE was of for ESA/ESTEC to measurements the AOCS Unit ulation the MACS attitude ACC via En demonstrate the new approach with re-usable hardwarethe stress distribution in aand vironment Simlation control databus. Dynamics ACC joint configuration. be In the lap the ureceived data will performed, using an analytical solution for T into the attitude control laws, which results in est Equipment and software [Ref. the results of an experimental programme on similar strip specimens is given. fed In reference 3 2]. This prototype has recently been developed into a fullblown AOCS test system commanding of the actuator units. of Here, the influence of various manufacturing techniquesestand stiffener flange The response thethe 3 Architectur o Fig. 2 units igur measured with a details on device Fig. e able to meet the requirements for both subsystem and actuator T confis ation monitoring strength of the skin-stiffener interface was established. routedspecimens were tested under lateral The back to the corresponding dynamics and system level testing of the AOCS of the XMM and and tension satellites. INTEGRALof the skin and with a pull-off test. environment model. In this way the loop is closed. Research using a similar type of specimen is reported in references 4 and 5. However,2. TEST CONCEPT The MACS interface has to the specimens used total) further simplified by eliminating the stiffener web. be programmable to (9 pages in were Figurereference a4 schematic overviewthe taper on the flange was combination ofunder three-pointunits. If 1 gives of a generic reflect any investigated real and simulated and In the influence of AOCS for spacecraft. The diagram reflects the cyclic real sensor and/or actuator units are not available they

four-point bending configurations. The results were evaluated using finite element analyses. In reference 5 four different skin laminates and two flange laminates were used to establish the

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TEST AND VERIFICATION EQUIPMENT FOR THE ATTITUDE & ORBIT CONTROL SYSTEM OF THE XMM SATELLITE Contents by

influence of skin and flange lay-up on Schenau, L.C.J. van Rijn, J. Spaa H.A. van Ingen delamination initiation. Results were again evaluated 3 Abstract using finite element analyses. National Aerospace on-going research performed at the National Laboratory NLR The present paper describes recent and P.O. which AD Emmeloord, of Netherlands Aerospace Laboratory NLR, Box 153, 8300at the derivationThe a failure criterion for stiffener pop3 1. Introduction Tel. +31 aims 218; Fax: +31 527 248 210 527 248 off and the implementation of this failure criterion in PANOPT. E-mail: ingen@nlr.nl; lvrijn@nlr.nl; spaa@nlr.nl Firstly, the test configuration used in the current work and the experimental results in terms of 3 failure loads Tensor Skin Concept ABSTRACT 2. and failure modes, is given. Then, the derivation of a A complete AOCS, based on nature of the AOCS. failure criterion together with dynamics this criterion can be considered as these results is presented. Subsequently, an evaluation ofand environment using experimental a The National Aerospace Laboratory isNLR in Finally, conclusions is actively closedfor future workControl the loop which and directions by the Attitude are results from references 4 and 5 given. 4 3. Experimental Results Netherlands has developed a new generation of Test Computer (ACC). given.and Verification Equipment (TVE) for testing of Attitude and 4. Simulations Orbit Control Subsystems of spacecraft. Based on a prototype TVE developed for ESA, test EXPERIMENTAL INVESTIGATION measurements equipment has been developed for Matra Marconi Space for AOCS subsystem performed Tests testing and system level and Future Developments 5. Recently of SPECIMEN INTEGRAL scientific satellites. the XMM and DESCRIPTIONSENSORS target attitude CONTR L O commands

7 9ACTU TORS A

9 This paper 6. Conclusions describes the test concept and the A simple test specimen configuration, similar to the configuration used in references 4 architecture of the XMM test system with its main DYNAMICS and 5, was used to development significant influence the strength ofmagnetic a skintorques features, the incrementaldetermine theand delivery, parameters whichattitude Info on: fields 9 References during development and use stiffener interface in and experiences obtained a composite stiffened panel. Earth disturbances Sun of the system. specimens consisted ofalsostiffener flange with a 45 taper, bonded onto a skin The The described work has a been Stars ENVIR ONMENT performed under ESA contract. width was equal to 50 mm. Specimen dimensions are given in figure 1, laminate. The specimen Magnetic field1. INTRODUCTION 4 experimental programme encompassed 5 Fig. 1 laminates which are combined with 4 ol system The Tables skin Generic attitude contr Based on experiences with the production and use of 8 Figures flange laminates, as given in tables I and II. various test systems for the ISO, SAX, SOHO and The laminates were made using Fibredux 6376C HTA prepreg, with a nominal thickness of 0.181 mm.the integration and test to bond the flange and other satellites, the National Aerospace Laboratory In The adhesive used phase the AOCS subsystem Physical Le vel skin the Netherlands 300M. NLR inlaminates was FMhas developed a new is gradually built up Attitude depending on the schedule of Data generation First, the Test and Verificationfor skins and flanges were manufactured and cured, except of generic laminated plates Equipment incoming units. Verification of attitude control modes Contr l Computer o AOCS (TVE) skin laminate S5. Flanges were then cut to the and real-time behaviour is done in the early period of with re-usable hardware and software for for required shape. Subsequently, MACS skins and the b us testing of Attitude and Orbit Control Subsystems integration using a combination of real and simulated flanges were co-bonded, except for skin S5. Skin S5 was assembled in a wet' condition with (AOCS) of spacecraft [Ref. 1]. units. Protocol Le vel

Syst

in which the fibre direction of a 0 ply (parallel to the stiffener axis) is also indicated.

Sy B Int

the cured flange F4. Finally, all specimens were cut to the indicated dimensions. SensorElectr nics o

The TVE had to be usable from the early stage of the AOCS development up to the integration of the AOCS in EXPERIMENTAL PROCEDURE tests with the spacecraft environment i.e. open loop a single unit up to closed loop tests with any combination of real and simulated AOCS units should The specimens were loaded in four-point be supported.

The test concept described in this paper is based on a static closed loop test facility (no real motion). The Sensor test configuration is shown in figureActuator dynamics 2. The Head User v and environment simulation is responsible for the Leel Data computation Electrical of stimuliCSfor the CS sensor Electrical and the units ing Physical bending in the sameMA unitrigMA used in earlier (Engineer test as units) processing of uli Stimmonitor ulation from the actuator units. Stim uli Simdata Monitoring Monitoring research (Ref. 2). A schematic representation of the loading configuration is will deliver sensor given in figure 2. The stimulated sensors Gr Load was applied for ESA/ESTEC to A prototype TVE was built under displacement control, with a constant AOCS Unit Sim rate ofMACS attitude measurements to the loading the 2 mm/min ACC ulation via En for the the new approach with re-usable (skins demonstratespecimens with thinner skinshardware S1 and S2) and with a constant Simlation control databus. Dynamics and vironment loading rate of 1 be In the ACC the ureceived data will T into the attitude control laws, which results in est Equipment and software for the2]. This prototype the thicker skins fed mm/min [Ref. specimens with has recently (skins S3, S4 and S5). been developed into a fullblown AOCS test deflection was monitored the actuator units. The response of the commanding of using two laser displacement During the test, the mid-span system Fig. 2 est igur Fig. 3 A rchitectur o e able to meet the requirements for both subsystem and actuator T conf ation transducers, which were plotted against the applied load. units is measured with athe cross head The displacement of monitoring device system level testing of the AOCS of the XMM and and routed back to the corresponding dynamics and was also monitored. Moreover, the specimen were monitored visually duringthe loop is using a INTEGRAL satellites. environment model. In this way loading closed.

Actuator Electr nics o

Data

photocamera which was zoomed in on the flange region. For several tests, acoustic emission monitoring equipment was also used. 2. TEST CONCEPT The MACS interface has to be programmable to (9 pages in total) Figure 1 gives a schematic overview of a set-up is shown inany combination of real and simulated units. If reflect figure 3. A picture of the experimental genericAOCS for spacecraft. The diagram reflects the cyclic real sensor and/or actuator units are not available they

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TEST AND VERIFICATION EQUIPMENT FOR THE ATTITUDE & ORBIT CONTROL SYSTEM OF THE XMM SATELLITE Contents by

EXPERIMENTAL RESULTS van Ingen Schenau, L.C.J. van Rijn, J. Spaa H.A.Abstract 3National Aerospace Laboratory for The delamination onset moments, which were foundNLRthe various specimens, are given P.O. Box 153, 8300 AD Emmeloord, The Netherlands table III. Introduction 3 1. Tel. +31 527 248 218; Fax: +31 527 248 210 It was established that theingen@nlr.nl; lvrijn@nlr.nl; spaa@nlr.nl skin laminate, exhibited a specimens, containing the same E-mail:

in

similar behaviour. Some observations for each set of specimens is given in this section. 3 Delamination could not be ABSTRACT 2. Tensor Skin Concept induced in the specimens with skin S1 using AOCS, together with nature of the AOCS. A complete the current test configuration. For these specimens the magnitudedynamicsdeflection was so can be that slippage a of the and environment large considered as The National The moment given in NLR in the the maximum moment applied by the the test Control Aerospace Laboratory table III is loop which is actively closed until Attitude was occurred. 3. Experimental Results 4 Netherlands has developed a new generation of Test Computer (ACC). stopped. and Verification Equipment (TVE) for testing of The load-deflection curveof spacecraft. for specimens containing skin S2 showed a decrease in load Attitude and 4. Simulations Orbit Control Subsystems 7 target could when a delamination was formed, ESA, the point of delamination onset attitude therefore be Based on a prototype TVE developed for and test determined easily. Non-linear behaviour measurements commands equipment has been developed for Matra Marconi was noted just before the formation of the CONTR L O Space for AOCS subsystem performed Tests testing delamination. Recently and system level and Future Developments 9 5. of Also, the and INTEGRAL scientific satellites. emission energy versus time showed a sharp increase the XMM graph of the cumulative acoustic SENSORS ACTU TORS A in energy, which corresponded to the formation of a delamination. 9 This paper 6. Conclusions describes the test concept and the The photographs taken during testing and post-mortem inspection of the specimens showed that architecture of the XMM test system with its main delaminations were formed between the 0 on DYNAMICS torques of features, the incremental development and delivery, and +45 layersattitude the right hand sidemagnetic the Info on: fields 9 References during development and layers on thethleft hand side of the specimen. A matrix specimen and between the +45 and -45 use and experiences obtained disturbances Ear of crack in the outermost 0 ply was observed at Sun the system. The described work has also been some stage before the formation of a Stars ENVIR ONMENT Syst performed under ESA photograph of a specimen containing delaminations on either side is given in contract. delamination. A Magnetic field figure 4. 1. INTRODUCTION 4 load-deflection curve for specimens containing skin S3 had an area of non-linear Fig. 1 Generic attitude contr ol system The Tables Sy Based on experiences with the production and use of B 8 Figureswas behaviour whichfor thequite large, SOHO and of the applied load occurred in most instances and a drop various test systems ISO, SAX, Int only at final failure, making it more difficult other satellites, the National Aerospace Laboratory to establish the point of delamination onset. In the integration and test phase the AOCS subsystem Physical vel The acoustic emission has developed not provide is clear-cut built up Attitudeof delamination onset of Le a gradually indication NLR in the Netherlands equipment did a new depending on the schedule Data generation of generic Testin the cumulative acoustic emission energy Verification of attitudethroughout incoming units. was Contr l Computer o either: the increase and Verification Equipment quite gradual control modes AOCS (TVE) loading processhardware and software for with re-usable upto the point of final failure.and real-time behaviour is done in the early period of the MA b CS us testing of Attitude and Orbit Control Subsystems integration using The photographs taken during testing and post-mortem inspection a combination of real and simulated of the specimens showed that (AOCS) of spacecraft [Ref. 1]. units. vel delaminations were formed in the adhesive layer on the right hand side of the specimen and Protocol Le Sensor Actuator Data Electr nics o Electr nics o between to +45 and 0 layers on of left The TVE hadthe be usable from the early stagethe the hand side of the described in Figure 5 is based on a The test concept specimen. this paper shows a delamination between integration of 0 layers AOCS development up to thethe +45 andthe AOCS on the left hand loop test a specimen. Substantial static closed side of facility (no real motion). The Sensor in growth of the delamination was seen in with the spacecraft environment i.e. open loop tests some instances. On the left hand side aActuator of the test configuration is shown in figure 2. The dynamics jump Head User v a delamination toto closed loop interface between the 0 environment simulation observed. single unit up the adjacent tests with any and and -45 layers was is responsible for the Leel Data combination of real and simulated AOCS units should computation Electrical of stimuliCSfor the CS sensor Electrical and the units ing Physical MA unit MA The results obtained for specimens with skin S4 were very similar to those obtained for (Engineer units) be supported. processing of uli Stimmonitor ulation from the actuator units. Stim uli Simdata Monitoring Monitoring specimens containing skin S2, albeit at considerably higher loads. sensors will deliver sensor The non-linearity prior to The stimulated Gr A delamination was absent for these specimens. prototype TVE was built for ESA/ESTEC to measurements to the AOCS Unit Sim the MACS attitude ACC ulation via En The delamination locations were the same demonstrate the new approach with re-usable hardware as found for specimens containing skin S2. The be control databus. Dynamics and vironment Simlation In the ACC the ureceived data will T into the attitude control laws, which results in est Equipment and software [Ref. delamination in specimen S4F3 is fed formation of a 2]. This prototype has recently shown in figure 6. been developed results fullblown specimens system The into a for the AOCS test containing commanding of the actuator units. The response of the 3 Architectur o skin 2S5 were again quite similar to those Fig. Fig. est igur measured e able to meet the requirements for both subsystem and actuator T confis ation units obtained for the specimens containing the skin S3. For this specimens the with a viable method only monitoring device system level testing of the AOCS of the XMM and and routed back to the corresponding dynamics and to determine the INTEGRAL satellites. delamination onset load was visual inspection, since the onset ofloop is closed. environment model. In this way the delamination was not evident on the load-deflection curve. Delamination was observed between the +45 and -45 layers on the left hand side of the to 2. TEST CONCEPT The MACS interface has to be programmable (9 pages in total) Figure 1 gives a schematic overview of a generic reflect any combination of real and simulated units. If specimen.AOCS for spacecraft. The diagram reflects the cyclic real sensor and/or actuator units are not available they

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TEST AND VERIFICATION EQUIPMENT FOR THE ATTITUDE & ORBIT CONTROL SYSTEM OF THE XMM SATELLITE Contents by

RESULTS FROM EARLIER INVESTIGATIONS van Rijn, J. Spaa H.A. van Ingen Schenau, L.C.J.Abstract 3National reported in reference 2 The results obtained earlier and Aerospace Laboratory NLR were again incorporated in the P.O. Box 153, 8300 were used earlier readily complement the parameters current work since the parameters which AD Emmeloord, The Netherlands 3 1. Introduction Tel. +31 527 248 218; Fax: +31 527 248 210 used in the current work. E-mail: ingen@nlr.nl; lvrijn@nlr.nl; spaa@nlr.nl

3 nature of the AOCS. A complete AOCS, together with dynamics and environment can be considered as a Specimen description The National Aerospace Laboratory NLR in the loop which is actively closed by the Attitude Control 4 3. Experimental Results Netherlands has specimensaused in reference 2Test developed new generation of consisted of a blade-type stiffener bonded onto a skin Computer (ACC). The and Verification Equipment (TVE) for testing of laminate. Orbit stiffener flange was not tapered. The specimen width was equal to 50 mm. The The Control Subsystems of spacecraft. Attitude and 4. Simulations 7 target overalladimensions of the specimens were the same as given in figure 1. attitude Based on prototype TVE developed for ESA, test three measurements equipment The been developed programmeMarconi has experimental for Matra encompassed 4 different skin lay-ups, commands different CONTR L O Space for AOCS subsystem performed Tests lay-up one or two variants with either a thicker stiffener and system level and Future Developments stiffener lay-ups and for each stiffener testing 9 5. Recently of blade, which INTEGRAL scientific satellites. the XMM and was obtained by adding a centre laminate to the blade, or a wider stiffener flange.ABSTRACT 2.

Tensor Skin Concept

This paper 6. Conclusions describes the test concept and the 0.181 mm. The adhesive used to bond the stiffener to the skin was FM 300K. architecture of the XMM test system with its main DYNAMICS The the incremental development and delivery, attitude torques features,skin and stiffener identification of all specimens is given in table IV. Info on: References during development and use and experiences obtained disturbances Earth Sun of the system. The described work has also been Stars ENVIR ONMENT performed under ESA contract. Experimental procedure Magnetic field

The laminates were made using Fibredux 6376/T400H prepreg, with a nominal thickness of9magnetic fields

SENSORS

ACTU TORS A

9

Syst

1. INTRODUCTION 4 same Fig. 1 Generic attitude contr system The Tables test set-up as shown in figure 2 was used to load theolspecimens in four-point Based on experiences with the production and use of bending. 8 Figures were performed under displacement control with a constant loading rate of The test various test systems for the ISO, SAX, SOHO and 1 satellites, the National Aerospace Laboratory othermm/min. Specimens were also tested under lateral tension, in whichtest phase theappliedsubsystem In the integration and the load is AOCS to the vel skin the Netherlands has developed a new control on a schedule Physical Le NLR in laminate only. The tests were performed under displacementAttitude is gradually built up dependingwith the constant of Data generation of generic1Test and Verification the tests, the applied load and theContr l Computerattitude control modes incoming units. Verification of o loading rate of mm/min. During Equipment cross head displacement (TVE) with re-usablePhotographs were taken for hardware and software during theAOCS and testing. behaviour is done in the early period of real-time were monitored. MA b CS us testing of Attitude and Orbit Control Subsystems integration using a combination of real and simulated (AOCS) of spacecraft [Ref. 1]. units. Protocol Le velSensor Electr nics o Actuator Electr nics o Data

Sy B Int

Experimental usable The TVE had to beresults from the early stage of the The test concept described in this paper is based on a AOCS development up to the integration of the AOCS static closed loop test facility (no real motion). The Sensor in the spacecraft environment i.e. open loop tests with test configuration is shown in figureActuator 2. The At a All specimens loaded in four-point bending showed a similar failure behaviour.dynamics Head User v a certain unit up to a crack loop tests in the any of the bond layer. After a further load increase,the Leel single load level closed initiated with fillet and environment simulation is responsible for Data combination of real and simulated AOCS units should computation Electrical of stimuliCSfor the CS sensor Electrical and the units ing MA unit MA a delamination formed either in the bond layer or inPhysical first layer of the skin laminate. The (Engineer the units) be supported. processing of uli Stimmonitor ulation from the actuator units. Stim uli Simdata Monitoring Monitoring bending moments at delamination initiation for all specimens are given in table V. The stimulated sensors will deliver sensor Gr The failure behaviour ESA/ESTEC to A prototype TVE was built for of the specimens loaded in lateral to the AOCS Unit Simspecimen for each measurements tension (oneulation the MACS attitude ACC via En configuration) approach with re-usable hardware demonstrate the newwas quite similar to that of the specimens loaded the and vironment Simlation control databus. Dynamics four-point bending. The be In in ACC the ureceived data will T into the attitude control laws, which results in est Equipment and softwaredelamination initiation for all specimens fed given in table VI. loads at [Ref. 2]. This prototype has recently are been developed into a fullblown AOCS test system commanding of the actuator units. The response of the Fig. 2 est igur measured with a monitoring device 3 A Fig. rchitectur o e able to meet the requirements for both subsystem and actuator T confis ation units system level testing of the AOCS of the XMM and and routed back to the corresponding dynamics and EVALUATION INTEGRAL satellites. OF EXPERIMENTAL RESULTS environment model. In this way the loop is closed. OBSERVATIONS AND INTERPRETATION OF EXPERIMENTAL has to be programmable to 2. TEST CONCEPT The MACS interface RESULTS (9 pages in total) Figure 1 gives a schematic overview of a generic reflect any combination of real and simulated units. If AOCS for The delamination onset moments for the specimens tested in the current programme as spacecraft. The diagram reflects the cyclic real sensor and/or actuator units are not available they given in table III are shown in figure 7. The delamination onset moments for the specimens

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TEST AND VERIFICATION EQUIPMENT FOR THE ATTITUDE & ORBIT CONTROL SYSTEM OF THE XMM SATELLITE Contents by

tested in four-point bending in van Ingen Schenau, L.C.J. van Rijn, J. in table V are shown in figure H.A. the earlier programme as given Spaa 3 Abstract 8. The delamination onset loads for the specimens tested in lateral tension in the earlier National Aerospace Laboratory NLR programme as given in table VI are shown in figure 9. In 1. Introduction Box 153, 8300 AD Emmeloord, The lay-up did not have a significant figure 7 it is P.O. that the variation in flange Netherlands seen 3 Tel. +31 527 248 218; Fax: +31 527 248 210 influence on the delamination onset moment.lvrijn@nlr.nl; spaa@nlr.nl This might be expected, since the delamination E-mail: ingen@nlr.nl; location is in the skin and the interaction between flange and skin is governed by the bending 3 stiffnesses the flanges which ABSTRACT 2. of Tensor Skin Concept are quite similar forofthe AOCS. A complete AOCS, together with nature the flange laminates in the current dynamics and environment can be considered as a programme. The National Aerospace same is observed for the delamination onset moments fromAttitude Control Laboratory NLR in the loop which is actively closed by the the earlier However, the 4 3. Experimental Results Netherlands has developed a new Here it is seen that for Computer (ACC). laminates 1a and 1b neither generation of Test research shown in figure 8. the thinner skin and Verification Equipment (TVE) for testing of variation Orbit Control lay-up nor of spacecraft. web thickness had a significant influence on the in stiffener Subsystems variation in Attitude and 4. Simulations 7 target attitude delamination onset moment. The ESA, test of influence of either stiffener lay-up or web Based on a prototype TVE developed for absence thickness is also developed for Matra Marconi it again measurements equipment has been observed for the skin laminate 2b. For skin laminate 2a OL was commands observed CONTR Space for AOCS subsystem performed Tests testing and system level and influence, but for this that the stiffener lay-up had no significant Future Developments laminate an increase in web 9 5. Recently of thicknessand INTEGRAL scientific satellites. the XMM (configuration 2a1c) or widening of the stiffener flange (configuration 2a3b) appeared SENSORS ACTU TORS A to have a decreasing influence on the delamination onset moment. 9 This paper 6. Conclusions describes the test concept and the The absence of an influence of the stiffener lay-up on the delamination onset moments is architecture of the XMM test system with its main DYNAMICS unexpected, especially in view of delivery, of magnetic attitude features, the incremental development and the variation by more than a factor 3 torquesthe bending Info on: fields 9 References during development and use stiffnesses obtained and experiencesof the stiffener laminates. disturbances Earth Sun of the system. The 9 it can be seen that neither variation in stiffener lay-up nor variation in the In figure described work has also been Stars ONMENT performed under ESA contract. web thickness had a significant influence on the delamination onsetENVIR for all skin laminates. load Magnetic field Observations of the delamination locations revealed that delamination always occurred 1. INTRODUCTION Fig. 1 Generic attitude contr ol system on either 4 Tablesthe 45 layer nearest to the flange. A schematic representation of the situation side of Based on experiences with the production and use of 8 Figures at delamination for the given in SOHO and various test systems onset isISO, SAX, figure 10, both for the configuration in which a 45 layer is situated adjacent to the Aerospace Laboratory other satellites, the Nationalflange and for the configuration in which a 0 test phasesituated between In the integration and layer is the AOCS subsystem Physical Le vel the first 45 layer and has developed a new NLR in the Netherlands the flange. is gradually built up Attitude depending on the schedule of Data generation findings are similar to the typical failure patterns described inContr l Computerattitude control modes incoming units. Verification of o These of generic Test and Verification Equipment reference 4.(TVE) with re-usable hardware and software for testing of Attitude and Orbit Control Subsystems (AOCS) of spacecraft [Ref. 1].AOCS and real-time behaviour is done in the early period of MA b CS us integration using a combination of real and simulated units. Protocol Le velSensor Electr nics o Actuator Electr nics o Data

Syst

Sy B Int

CRITERION DEFINITION

The TVE had to be usable from the early stage of the The test concept described in this paper is based on a Based up the integration of the AOCS AOCS development on to the above observations (no significant influence facility (no real motion). The static closed loop test of flange lay-up, and Sensor in delaminationenvironment occurring at the 45 layer nearest to the flange), it is postulated that the the spacecraft initiation i.e. open loop tests with test configuration is shown in figureActuator dynamics 2. The Head User v a average in-plane normal loop tests with any to the fibre direction and theisaverage in-planethe Leel single unit up to closed strain perpendicular and environment simulation responsible for Data combination of real and simulated AOCS units should computation Electrical of stimuliCSfor the CS sensor Electrical and the units (Engineer ing Physical MA unit MA shear angle in the 45 layer nearest to the flange govern the initiation of a delamination. units) be supported. processing of uli Stimmonitor ulation from the actuator units. Stim uli Simdata Monitoring Monitoring Obviously, the ratio between the normal strain perpendicular to the fibre direction sensor 22 The stimulated sensors will deliver Gr A and the shear angle built depends on the skin laminate. A laminate analysis programme LAP prototype TVE was 12 for ESA/ESTEC to measurements to the AOCS Unit Sim the MACS attitude ACC ulation via (Ref. 6) was used to obtain the normal strain demonstrate the new approach with re-usable hardware perpendicular to the the and vironment Simlation 22 data will be control databus. Dynamics ACC the ureceived and the In fibreEndirection T into the attitude control laws, which results in est Equipment and software [Ref. as functions of the applied moment or as functions of the applied lateral force. fed shear angle 12 2]. This prototype has recently been developed strains caused by the curing process, are also taken into account. The response of the into a fullblown AOCS test system commanding of the actuator units. The initial Fig. 2 est igur measured with a monitoring device 3 A Fig. rchitectur o e able to meet the requirements for both subsystem and actuator T confis ation units system level testing of the AOCS of the XMM and and routed back to the corresponding dynamics and For the evaluation of the experimental results, the following material the loop is closed. INTEGRAL satellites. environment model. In this way parameters were

used for the prepreg materials Fibredux 6376C HTA and Fibredux 6376/T400H: Ex = 137000 2. TEST CONCEPT MPa The MACS interface has to be programmable to (9 pages in total) Figure 1 gives a schematic overview of a generic reflect any combination of real and simulated units. If Ey = 9500 MPa AOCS for = 5100 MPa diagram reflects the cyclic spacecraft. The real sensor and/or actuator units are not available they Gxy xy = 0.3

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TEST AND VERIFICATION EQUIPMENT FOR THE ATTITUDE & ORBIT CONTROL SYSTEM OF THE XMM SATELLITE Contents

= -1 10-6 K-1 H.A. van Ingen Schenau, L.C.J. van Rijn, J. Spaa 3 Abstract = 3 10-5 K-1 National Aerospace Laboratoryto a room temperature of 20 C) was NLR A temperature difference of -150 C (from 170 C P.O. the initial strains. used in the determination ofBox 153, 8300 AD Emmeloord, The Netherlands 3 1. Introduction Tel. +31 527 248 218; Fax: +31 527 248 210 For the specimens with skin laminate S2 and S4 the presence of the matrix crack in the E-mail: ingen@nlr.nl; lvrijn@nlr.nl; spaa@nlr.nl outermost 0 layer was taken into account by using the following material properties for the 3 outermost layer: ABSTRACT 2.0 Tensor Skin Concept nature of the AOCS. A complete AOCS, together with dynamics and environment can be considered as a Ex = 137000 MPa The National1.10-5 MPa Laboratory NLR in the Aerospace loop which is actively closed by the Attitude Control = 3. Experimental Results Ey 4 -5 Netherlands has developed a new generation of Test Computer (ACC). Gxy = 1.10 MPa and Verification Equipment (TVE) for testing of xy =0 Attitude and 4. Simulations Orbit Control Subsystems of spacecraft. 7 target attitude Thereby, TVE developed capable test Based on a prototypethis layer is not for ESA, of carrying any load but partly restrains the Poisson's contraction been developed for Matra Marconi measurements commands equipment has of the laminate. CONTR L O Space for AOCS subsystemthe tests onlevel testing withDevelopments laminates rendered a failure and system specimens Evaluation of performed Tests and Future various skin 9 5. Recently of envelope and the the XMM in INTEGRAL scientificAs an example, ellipses are used to fit to the current data sets: satellites. 22-12 plane.SENSORS ACTU TORS A

x y

by

This paper 6. Conclusions describes the test concept and 2 the architecture of the XMM test system with its main 22 f features, the incremental development and delivery, References and experiences obtained during development and use 22 of the system. The described work has also been performed under ESA contract.

12 f 12

1 (1) attitude Info on: Earth Sun Stars Magnetic field

2

9DYNAMICS disturbances ENVIR ONMENT torques magnetic fields

9

Syst

1. INTRODUCTION Tables Fig. 1 Generic attitude ol system figure 11.4The failure strains used to compose the criterion were 0.5 contr f22 and 1.75 for f12. for Based on experiences with the production and use of 8 seen, the variation in As can be Figures the ISO, SAX,the ratioand 22 and 12 is only minor for the skin laminates of various test systems for SOHO S1, S2, S3 the National Aerospace Laboratory other satellites, and S4. The ratio is approximately equalthe integration and test phase andAOCS subsystem In to 0.4. The ratio of 22 the 12 for skin Physical Le vel S5 in the Netherlands has developed a new NLR is approximately equal to 0.25. is gradually built up Attitude depending on the schedule of Data generation of generic Test and Verificationresults for specimens with the thinnerComputerlaminate and modes incoming units. VerificationS2 attitude control the of Contr l o The criterion brings together the Equipment AOCS (TVE) with re-usable hardware and software S3, S4 and real-time behaviourThe plotted strains for of for is done in the early period results for specimens with the thicker and S5 laminates. MA b CS us testing of Attitude and Orbit Control Subsystems integration using a combination of real and simulated specimens with skin S1 pertain to the maximum loads during the tests, which were not (AOCS) of spacecraft [Ref. 1]. units. Protocol Le vel

The failure envelope for the specimens tested in the current programme is given in

Sy B Int

sufficient to cause delamination. Clearly, the results obtained for skin S1 are not in line with Data Sensor Actuator Electr nics o Electr nics o the proposed criterion. The TVE had to be usable from the early stage of the The test concept described in this paper is based on a The failure the integration the AOCS AOCS development up to envelope for of the specimens testedclosed loop test facility (no in the earlier static in four-point bending real motion). The in programme is given in i.e. open12. The failure strains configuration is shown in criterion The dynamics the spacecraft environment figure loop tests with test Sensor to compose the figureActuatorfor these 2. used Head User v a experimental results wereloop tests with any for f which are quite similar to the values usedthe Leel single unit up to closed 0.35 for f and 2 and environment simulation is responsible for Data 22 12 combination of real and simulated AOCS units should computation Electrical of stimuliCSfor the CS sensor Electrical and the units ing Physical MA unit in figure 13. It should be kept in mind that the flanges of the specimens MA with which these (Engineer units) be supported. processing of uli Stimmonitor ulation from the actuator units. Stim uli Simdata Monitoring Monitoring experimental results were obtained, were not tapered. stimulated sensors will deliver sensor The Gr for skin laminates AOCS Unit Sim2a, and for 1b and A As can be TVE was ratio of 22 and 12 is similar measurements to the 1a andulation the MACS attitude prototype seen, the built for ESA/ESTEC to ACC via En 2b respectively. The sign of the shear angle, since the the received data will demonstrate the new approach with re-usable hardware which was negativeDynamics and vironment Simlation had a -45 be control databus. In the ACCtop ulayer T into the attitude control laws, which results in est Equipment and software [Ref. 2]. This prototype has recently into the first quadrant. fed orientation, was changed to bring all results been developed into afigure 12 the criterion appears to commanding of the actuator units.specimens withthe fullblown AOCS test system For the results in beFig. 2 est igurthe results for The response of Fig. 3 Architectur o applicable: e able to meet the requirements for both subsystem and actuator T confis ation units measured with a monitoring thicker and thinner skin laminates, skin laminates 2a and 2b, and skin laminates 1a and device 1b system level testing of the AOCS of the XMM and and routed back to the corresponding dynamics and respectively, are results for very different the loop is can be INTEGRAL satellites. equally well described; also, theenvironment model. In this way laminates closed. described by the proposed criterion. It should be noted 2. TEST CONCEPT that the skin laminate 1b from the earlier programme is identical to the skin to The MACS interface has to be programmable (9 pages in total) Figure 1 gives a in the current programme. The same applies for skinoflaminates 2b and S3 If reflect any combination real and simulated units. laminate S1 schematic overview of a generic AOCS for spacecraft. Thethe fact that the criterion appears to properly describe the behaviour of diagram reflects the cyclic real sensor and/or actuator units are not available they respectively. From specimens containing skin laminates 1b, whereas it appears to be less appropriate for specimens

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containing skin laminate S1, it might be Schenau, L.C.J. van Rijn, J. Spaa of the test set-up under H.A. van Ingen concluded that the performance 3 Abstract large deflections should be evaluated. National Aerospace Laboratory NLR The failure envelope for the specimens tested in lateral tension in the earlier programme P.O. Box 153, The Netherlands is given in figure 13. The ratios of8300 AD Emmeloord,slightly different under lateral tension 3 1. Introduction Tel. +31 527 22 and 12 are 527 248 210 248 218; Fax: +31 compared to the ratios under bending, but the same observations as given above, apply. The E-mail: ingen@nlr.nl; lvrijn@nlr.nl; spaa@nlr.nl failure strains used to compose the criterion for these experimental results were 0.3 for f22 and 3 1.3 for f2. Tensor Skin Concept ABSTRACT12. Again these values are quite similar to the of the AOCS. A complete 12. Clearly, the nature values used in figure AOCS, together with dynamics and environment presently not clear mode of loading should not be an important parameter in the criterion. It iscan be considered as a The National Aerospace Laboratory NLR results pertainingwhich is actively tension loading and Control in the loop to the lateral closed by the Attitude the whether the difference between the 4 3. Experimental Results Netherlandspertaining to bending should be considered statistically significant. has developed a new generation of Test Computer (ACC). resultsand Verification Equipment (TVE) for testing of Attitude and 4. Simulations Orbit Control Subsystems of spacecraft. Based on a prototype TVE developed for ESA, test CONCLUSIONS measurements equipment has been developed for Matra Marconi Space for AOCS subsystem performed Tests testing and system level and Future Developments 5. Recently of the XMM andderived failure criteria and the failure strains as given The INTEGRAL scientific satellites.target attitude CONTR L O commands

7 9

in the previous section were SENSORS ACTU TORS A used to calculate the delamination onset moments for the various skin laminates. These 9 This paper 6. Conclusions describes the test concept and the calculated delamination onset moments are plotted against the experimental delamination onset architecture of the XMM test system with its main DYNAMICS moments in figures development and delivery, current magnetic the and attitude features, the incremental 14 and 15 for the four point bending test results of the torques Info on: 9 References during development and use The black line in these figures indicates afields earlier experimental perfect and experiences obtained programme, respectively. disturbances Earth Sun of match between calculated work experimentally determined moments, both purple lines indicate the system. The described and has also been Stars ENVIR ONMENT performed underof 20contract. for the experimental results. field ESA percent a variation Magnetic If a margin of 20 percent is considered acceptable, it can be concluded that the results 1. INTRODUCTION 4 Tables Fig. 1 Generic attitude ol system of the current experimental programme are described quite well, as contr be seen in figure 18. can Based on experiences with the production and use of 8 Figures The calculated moment ISO,specimens with skin laminate S1 was found to be too conservative. for SAX, SOHO and various test systems for the The calculated National Aerospace Laboratory other satellites, the moment for specimens with skin laminate S2 was and test phase the AOCS subsystem In the integration almost exact. The calculated vel moment for specimens with skin laminate S3 conservative. However, Physical Le NLR in the Netherlands has developed a new wasisapparentlybuilt up Attitude gradually slightlydepending on the schedule of Data generation of be keptTestmindVerification Equipmentlower incoming units. detectability of a delamination, Contr l Computer o it should generic in and that in view of the than average Verification of attitude control modes (TVE) with re-usable hardware fromsoftware for and the large variation,AOCS experimental results could be slightly of and the real-time behaviour is done in the early period which may also be deduced MA b CS us testing of Attitude and Orbit Control Subsystems integration using a combination of real and biased. The calculated moment for specimens with skin laminate S4 was somewhat simulated high, (AOCS) of spacecraft [Ref. 1]. units. vel resulting in a non-conservative prediction. The variation in experimental resultsActuator again Protocol Le was Sensor Data Electr nics o Electr nics o substantial. be usable from the early stage of the The TVE had to The calculated moment for specimens with skin laminate S5 corresponded quite a The test concept described in this paper is based on well with the up to the integration of AOCS developmentexperimental results. the AOCS static closed loop test facility (no real motion). The Sensor in the spacecraft environment i.e. open loop tests with test configuration is shown in figureActuator dynamics 2. The From the comparison of the calculated and the experimental results form the earlier Head User v a experimental programme (Fig. 15), it can be concluded environmentresults are generally describedthe Leel single unit up to closed loop tests with any and that these simulation is responsible for Data combination of real and simulated AOCS units should computation Electrical of stimuliCSfor the CS sensor Electrical and the units ing Physical MA unit MA quite well. The extreme values mostly pertain to specimens with a stiffener which had a thicker (Engineer units) be supported. processing of uli Stimmonitor ulation from the actuator units. Stim uli Simdata Monitoring Monitoring stiffener web or had an enlarged stiffener flange, especially as far as the specimens with a The stimulated sensors will deliver sensor A thicker skin were concerned. ESA/ESTEC to prototype TVE was built for measurements to the AOCS Unit Sim the MACS attitude ACC ulation via

Syst

Sy B Int

Gr

En demonstrate the new approach with re-usable hardware control databus. Dynamics and vironment Simlation In the ACC the ureceived data will be T into the attitude control laws, which results in est Equipment and software [Ref. 2]. This prototype has recently fed been developed into a fullblown AOCSFROM THE LITERATURE the actuator units. The response of the test system commanding of EXPERIMENTAL RESULTS Fig. 2 est igur measured with a monitoring device 3 A Fig. rchitectur o e able to meet the requirements for both subsystem and actuator T confis ation units system level testing of the AOCS of the XMM and and routed back to the corresponding dynamics and The applicability of the criterion was also environment for experimentalthe loop is closed. evaluated model. In this way results given in INTEGRAL satellites.

references 4 and 5. In reference 4 results are given for a 24-ply skin interfacea has to be programmable to quasi-isotropic lay-up. 2. TEST CONCEPT The MACS with (9 pages in total) Figure 1 gives were similaroverview specimens used in the current experimental and simulated units. If reflect any combination of real programme. The Specimens a schematic to the of a generic AOCS for spacecraft. was equal to reflects the cyclic real manufactured from the are not available they specimen width The diagram 25 mm. Specimens weresensor and/or actuator units prepreg material IM6/3501-6. They contained either square ended (specimen code B) or tapered flanges

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(specimen code T). The experimental programme encompassed Spaa H.A. van Ingen Schenau, L.C.J. van Rijn, J. three-point bending tests at a Abstract 3 inch span, three-point bending tests at a 4 inch span and four-point bending tests. 3 National Aerospace Laboratory NLR For the evaluation of the experimental results the following material parameters were P.O. Box 153, 8300 AD Emmeloord, The Netherlands used for the prepreg material IM6/3501-6: 3 1. Introduction Tel. +31 527 248 218; Fax: +31 527 248 210 = 145000 MPa E-mail: ingen@nlr.nl; lvrijn@nlr.nl; spaa@nlr.nl Ex = 9600 MPa Ey 3 2. Tensor Skin Gxy ABSTRACT = 5200 MPa Concept nature of the AOCS. A complete AOCS, together with dynamics and environment can be considered as a xy = 0.3 -1 The National= Aerospace -7Laboratory NLR in the loop which is actively closed by the Attitude Control x 4 3. -3.6 10 K Results Experimental -5 new Netherlands has developed a -1 generation of Test Computer (ACC). y = 2.9 10 K and Verification Equipment (TVE) for testing of tlayer = 0.188 mm Attitude and 4. Simulations Orbit Control Subsystems of spacecraft. 7 target attitude Aprototype TVE difference for -150 C (from 170 C to a room temperature of 20 C) was Based on a temperature developed of ESA, test used in the been developed for Matra Marconi measurements commands equipment has determination of the initial strains. CONTR L O Space for AOCS subsystem performed Testsended flange were unadvertently manufactured with a The specimens and system level testing with a square and Future Developments 9 5. Recently of slightly different lay-up. The difference in lay-up is taken into account in the determination of the XMM and INTEGRAL scientific satellites. SENSORS ACTU TORS A the strains. The sign of the shear angle was changed to bring all results into the first quadrant. 9 This paper 6. Conclusions describes the test concept and the In figure 16 the calculated strains at failure and 12 are given. The most salient observation architecture of the XMM test system with its main 22 DYNAMICS is the considerably development at delivery, to the strains at magnetic attitude features, the incremental lower strains and failure for these specimens in comparison torques on: fields References during development strains are less Infoth 50 percent of the strains given 9 failure given in figures 11 to 13: the and use than there. and experiences obtained disturbances Ear Sun of As can be seen, the ratio of 22has also been the system. The described work and 12 is quite different for the skin laminate combined with Stars ENVIR ONMENT Syst performed under ESA contract. and the skin laminate combinedfield the square ended flanges Magnetic with the tapered flanges. As a result of the low mechanical strain levels, the relative influence of the curing strains is larger. 1. INTRODUCTION 4 Tables Fig. 1 Generic attitude contr ol system For the tapered specimens the strains at failure are similar for all three loading conditions. For Sy Based on experiences with the production and use of B 8 ended the square Figuresspecimens SAX, SOHOat failure obtained using the three-point bending tests the strains and various test systems for the ISO, Int are somewhat lower than those obtained using four-point bending test. The variation in the other satellites, the National Aerospace Laboratory the In the integration and test phase the AOCS subsystem Physical Le vel results the Netherlands has developed a new NLR in for specimens with square ended flanges is considerably lessup Attitude the specimens with of is gradually built than for depending on the schedule Data generation of generic Test and Verification presented here incoming units. Verification of attitude controlof a Contr tapered flanges. From the results Equipment it can be concluded ol Computer influence modes that the AOCS (TVE) with re-usable the delamination onset for absent real-time behaviour is done in the flange and of hardware and software was and for specimens with a tapered early period transverse load on MA b CS us testing of Attitude and Orbit Control Subsystems integration relatively small for specimens with a square ended flange. using a combination of real and simulated (AOCS) of spacecraft [Ref. 1]. units. vel Since there is only one point per configuration, it is not possible to accurately construct failure Protocol Le Sensor Actuator Data o Electr envelopes. From the from the early stage appears likelyElectr nicsthe applicable in this paper onics would a that The TVE had to be usable available results itof the The test concept described failure strainsbased on is be development up ended than for tapered specimens. The factloop test facility (no real motion). The AOCSless for square to the integration of the AOCS static closed that the flange tapering appears Sensor in to influence environment i.e. open onsettests with whereas the flangeis lay-up in figureActuator have no the spacecraft the delamination loop moment test configuration shown appears 2. The dynamics to Head User v a significant influence, would merit furtherany single unit up to closed loop tests with research. and environment simulation is responsible for the Leel Data combination of real and simulated AOCS units should computation Electrical of stimuliCSfor the CS sensor Electrical and the units ing Physical unit MA In reference 5, results are given of four-point bending tests.MAFive different specimen (Engineer units) be supported. processing of uli Stimmonitor ulation from the actuator units. Stim uli Simdata Monitoring Monitoring configurations were tested: four different skin laminates stimulated and one skin laminate with The (A,C,D,F), sensors will deliver sensor Gr A two different flange built for (A, B). All flanges were tapered with aAOCS Unittaper.the MACS attitude prototype TVE was lay-ups ESA/ESTEC to measurements to the 20 Sim ACC ulation The specimen via En width was new approach with re-usable hardware demonstrate the 25 mm. control databus. Dynamics and vironment Simlation In the ACC the ureceived data will be T into the attitude control laws, which results in est Equipment and software [Ref. 2]. was reported has have occurred in the flange for all configurations but fed The delamination This prototype to recently been developed into a Therefore the proposed criterion can not the actuator units. The response of the fullblown AOCS test system commanding of be considered valid, for all configuration D. Fig. 2 est igur Fig. 3 A rchitectur o e able to meet the requirements for both subsystem and actuator T confis ation units configurations but configuration D. For those configurations, themeasured at failure, which device strains with a monitoring are system level testing of the AOCS of the XMM and and routed back to the corresponding dynamics and shown in figure INTEGRAL satellites.17, should be viewed as lying inside the proposed failure envelope. Theclosed. environment model. In this way the loop is strains at failure were calculated using the same material data as used for the results of reference 4. Comparing the 2. TEST CONCEPTrest for configurations A and B, the influence of the flange to be programmable to The MACS interface has laminate lay-up on (9 pages in total) Figure delamination onset moment is again found toreflect any combination forreal and simulated units. If the 1 gives a schematic overview of a generic be negligible, also of delamination in the AOCS for spacecraft. The diagramthe outermost 45 layer for configuration C isunits aresmaller, since reflects the cyclic real sensor and/or actuator much not available they flange. The failure strain in this layer was the second layer from the flange interface while the layer adjacent to the flange

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interface had an orientation perpendicular to the flange Rijn, J. (a 90 layer in the coordinate H.A. van Ingen Schenau, L.C.J. van edge Spaa 3 Abstract frame given in figure 1). National Aerospace Laboratory NLR those based on the results of The strains at failure were approximately 3 times higher than P.O. Box 153, 8300 AD Emmeloord, The Netherlands reference 1. ItIntroductionnoted that the average delamination onset moment obtained by four 4. should be 3 Tel. +31 527 248 218; Fax: +31 527 248 210 point bending as given in reference 4 was 645 lvrijn@nlr.nl; while the average delamination onset Nmm/mm, spaa@nlr.nl E-mail: ingen@nlr.nl; moment obtained by four point bending as given in reference 5 was 790 Nmm/mm. The 3 relevant 2. Tensor Skin of the ABSTRACT bending stiffnes Concept 24-ply quasi-isotropic the AOCS. A complete AOCS, together with nature of skin laminate in reference 4 was dynamics and environment can be considered as approximately 120 Nm for all skin laminates. The difference in delamination onset moments a The National Aerospace4 Laboratory NLR line with the loop which is actively closed by the Attitude Control in the between references and 5 Resultsin experimental results presented in this paper. 4 3. Experimental is notNetherlands has developed a new generation of Test Computer (ACC). and Verification Equipment (TVE) for testing of Attitude and 4. Simulations Orbit Control Subsystems of spacecraft. CONCLUSIONS AND RECOMMENDATIONS Based on a prototype TVE developed for ESA, test measurements equipment has been developed for Matra Marconi Space for AOCS subsystem performed Tests testing and system level and Future Developments CONCLUSIONS 5. Recently of the XMM and INTEGRAL scientific satellites.SENSORS target attitude CONTR L O commands

7 9ACTU TORS A

9 This paper 6. Conclusions describes the test concept and the significant influence on the delamination onset moment in four point bending tests or the architecture of the XMM test system with its main DYNAMICS delamination load in lateral tension delivery, magnetic attitude torques features, the incremental development and tests. Info on: fields 9 References during development and locations revealed that delamination always occurred Observations and experiences obtained of the delamination use disturbances Earth of on either side of described layer nearest to the flange.Sun the system. The the 45 work has also been ENVIR ONMENT performed under proposed criterion brought together the Stars field for specimens with the thinner S2 The ESA contract. results Magnetic1. INTRODUCTION 4 Tables Fig. 1 Generic attitude ol system obtained for skin S1 were not in line with the proposed criterion. contr Based on experiences with the production and use of 8 criterion The Figures appeared to SOHO applicable for the results of the earlier experimental various test systems for the ISO, SAX, be also and programme. The results Aerospace Laboratory and thinner skin laminates were equally well other satellites, the National for specimens with thicker In the integration and test phase the AOCS subsystem Physical Le vel described. Netherlands has developed a new NLR in the Moreover, the results for very different laminates could be described by the proposed of is gradually built up Attitude depending on the schedule Data generation of generic Test and Verification Equipment incoming units. Verification of attitude control modes Contr l Computer o criterion. AOCS (TVE) with re-usable hardware in behaviour observed between the the skin laminate 1b and the skin of and software for and real-time behaviour is done in the early period From the difference MA b CS us testing of Attitude and Orbit Control Subsystems integration laminate S1, it was concluded that the performance of the usingset-up under large deflections test a combination of real and simulated (AOCS) of spacecraft [Ref. 1]. units. Protocol Le vel

For the skin laminates considered, a variation of the flange lay-up did not have a

Syst

laminate and the results for specimens with the thicker S3, S4 and S5 laminates. The results

Sy B Int

should be evaluated. Sensor Actuator Data Electr nics o Electr nics o The be usable from the early stage be an The TVE had tomode of loading should not of the important parameter in the criterion. It is unclear a The test concept described in this paper is based on whether the difference between the results AOCS development up to the integration of the AOCS pertaining closed loop test tension (no real motion). The static to the lateral facility loading and the Sensor in results pertaining to bending shouldtests considered statistically significant. figureActuator dynamics the spacecraft environment i.e. open loop be with test configuration is shown in 2. The Head User v a single unit upthe comparison of experimentally determined and calculated delamination onsetthe Leel to closed loop tests with any and environment simulation is responsible for From Data combination of real and simulated AOCS units should computation Electrical of stimuliCSfor the CS sensor Electrical and the units ing Physical MA moments, it can be concluded that the results of the current and MA unitearlier experimental (Engineer the units) be supported. processing of uli Stimmonitor ulation from the actuator units. Stim uli Simdata Monitoring Monitoring programme can be described quite well. The stimulated sensors will deliver sensor Gr The fact that the for ESA/ESTEC to A prototype TVE was built flange tapering appears to influence the the AOCS Unit Sim the MACS attitude measurements to delamination onset moment ACC ulation via En whereas the flange lay-up appears to have no demonstrate the new approach with re-usable hardware significantdatabus. Dynamics and vironment Simlation research. be control influence, merits the ureceived data will In the ACC further T into the attitude control laws, which results in est Equipment and software the specimen width would be an important parameter, this might restrict the relevance fed If [Ref. 2]. This prototype has recently been developed into a fullblownstrip specimens to the behaviour of panels. AOCS test system commanding of the actuator units. The response of the of the results obtained on Fig. 2 T conf ation est igur Fig. 3 A rchitectur o eable to meet the requirements for both subsystem and system level testing of the AOCS of the XMM and RECOMMENDATIONS INTEGRAL satellites. actuator units is measured with a monitoring device and routed back to the corresponding dynamics and environment model. In this way the loop is closed.

The applicability of the proposed criterion The MACS further assessed by the execution to should be interface has to be programmable 2. TEST CONCEPT (9 pages in total) Figure dedicateda experimental programme in which reflect any combinationlaminates resulting units. If of 1 gives schematic overview of a generic a number of skin of real and simulated in a AOCS for spacecraft.ratiosdiagram reflects the cyclic The real sensor and/or actuator units are not available they variety of strain The should be incorporated. programme should also encompass a number of flanges with various stiffnesses and tapers. Moreover, the influence of the specimen width

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on the delamination onset moment should be L.C.J. van Rijn, J. Spaaissue of testing thin skin H.A. van Ingen Schenau, investigated. The 3 Abstract laminates, which would generally be loaded upto large deflections, should also be addressed. National Aerospace Laboratory current simple test configuration and NLR A correlation between the results obtained with the P.O. Box to 8300 AD Emmeloord, The Netherlands results from Introductionneeds 153,be established. Possibly, bending tests on panels 3 an of 1. panel tests Tel. +31 527 248 218; Fax: +31 527 248 210 intermediate size could fillE-mail: ingen@nlr.nl; lvrijn@nlr.nl; spaa@nlr.nl the existing gap. Finally, when the criterion is proven to be applicable, it will be implemented in the 3 optimization Tensor Skin Concept ABSTRACT 2. code PANOPT, so the influence of the skin-stiffener interface strength on optimal nature of the AOCS. A complete AOCS, together with dynamics and environment can be considered as a panel designs can be determined.The National Aerospace Laboratory NLR in the loop which is actively closed by the Attitude Control 4 3. Experimental Results Netherlands has developed a new generation of Test Computer (ACC). and Verification Equipment (TVE) for testing of ACKNOWLEDGEMENTS of spacecraft. Attitude and 4. Simulations Orbit Control Subsystems 7 target attitude Based on a prototype TVE developed for ESA, test measurements commands equipment This been developed has been carried out under a contract awarded by the Netherlands has investigation for Matra Marconi CONTR L O Space for AOCSAerospace performed Tests testing number 01310 N. Agency for subsystem and system levelcontract Programmes, and Future Developments 9 5. Recently of the XMM and INTEGRAL scientific programme of a student from Imperial College, ms. S. Handley, satellites. The contributions to the This paper 6. Conclusions describes the test concept and the architecture of the XMM test system with its main features, the incremental development and delivery, References REFERENCES and experiences obtained during development and use of the system. The described work has also been performed under ESA contract. 1. P. Arendsen, H.G.S.J. Thuis, J.F.M.

and her supervisor K. Stevens, are gratefully acknowledged.Info on: Earth Sun Stars Magnetic field

SENSORS

ACTU TORS A

9attitude DYNAMICS disturbances torques magnetic fields

9

ENVIR ONMENT Syst Wiggenraad Optimization of composite stiffended panels with postbuckling constraints, National 1. INTRODUCTION 4 Tables ol Aerospace Laboratory Technical Publication Fig. 1 Generic attitude contrpublished in the Proc. NLR TP 94083 U,system Sy Based on experiences with the production and use of th B Forum the ISO, SAX, SOHO and 508 Figures of the American Helicopter Society, Washington, DC,1994 various test systems for Int 2. H.G.S.J. Thuis, Aerospace Laboratory other satellites, the National J.F.M. Wiggenraad In the integration and test phase the AOCS subsystem Physical Le vel NLR in the Netherlands ofhas developed a new of a isdiscrete skin-stiffener interface, National of gradually built up Attitude depending on the schedule Data Investigation the bond strength generation Aerospace Laboratory Technical Publication NLR TP 92183 Contr l1992 attitude control modes of generic Test and Verification Equipment incoming units. Verification of L,o Computer AOCS (TVE) with re-usable hardware and software for and real-time behaviour is done in the early period of 3. H.G.S.J. Thuis MA b CS us testing of Attitude and Orbit Control Subsystems integration using a combination of real and simulated Onvestigation 1]. the strength of bonded joints between composite stiffeneres and skins into (AOCS) of spacecraft [Ref. units. vel - continuation of the investigation described inSensor TP 92183 L -, National Aerospace Protocol Le NLR Actuator Data Electr nics o Electr nics o Laboratory Contract early stage of the The TVE had to be usable from theReport NLR CR 93279 Ctest concept described in this paper is based on a The (in Dutch), 1993 AOCS development up to the integration of the AOCS static closed loop test facility (no real motion). The 4. P.J. Minguet, T.K. O'Brien Sensor in the spacecraft environment i.e. open loop tests with test configuration is shown in figure 2. The Analysis of test methods for characterizing skin/stringer debondingActuator dynamics failures in Head a single unit up to composite panels, Composite materials: testing and design is responsible for the Leel and environment simulation (Twelfth volume), User v reinforced closed loop tests with any Data combination of real and simulated AOCS units should computation Electrical of stimuliCSfor the sensor Electrical and the units ing unit MA for Testing ASTM STP 1274, R.B. Deo and C.R. Saff, Physical AmericanMASociety CS the actuatorand (Engineer Eds., of uli units) be supported. processing Stimmonitor ulation from Stim uli Simdata Monitoring Monitoring units. Materials, 1994 The stimulated sensors will deliver sensor Gr P.J. Minguet, T.K. O'Brien A 5. prototype TVE was built for ESA/ESTEC to measurements to the AOCS Unit Sim the MACS attitude ACC ulation via En demonstrate the new approach with re-usable hardware bond failure using athe and vironment Simlation release rate be control databus. Dynamics ACC energy In strain the ureceived data will Analysis of composite skin/stringer T into the attitude control laws, which results in est Equipment and software [Ref. 2]. Proceedings ofhas recently Whistler, B.C., 1995 fed approach, This prototype ICCM-10, been developed into aAnalysis Program, User Guide, Anaglyph Ltd, London 1996 The response of the fullblown AOCS test system commanding of the actuator units. 6. Laminate Fig. 2 T conf ation est igur Fig. 3 A rchitectur o e

able to meet the requirements for both subsystem and system level testing of the AOCS of the XMM and INTEGRAL satellites.

actuator units is measured with a monitoring device and routed back to the corresponding dynamics and environment model. In this way the loop is closed. The MACS interface has to be programmable to reflect any combination of real and simulated units. If real sensor and/or actuator units are not available they

2. TEST CONCEPT (9 pages in total) Figure 1 gives a schematic overview of a generic AOCS for spacecraft. The diagram reflects the cyclic

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TEST AND VERIFICATION EQUIPMENT FOR THE ATTITUDE & ORBIT CONTROL SYSTEM OF THE XMM SATELLITE Contents by

TABLE I LAMINATE LAY-UPL.C.J. van Rijn, J. FOR SKIN (S) H.A. van Ingen Schenau, DEFINITION Spaa Abstract AND STIFFENER FLANGE SIMULATION (F) Code S1 F11.National Aerospace Laboratory NLR Laminate lay-up 8300 AD Emmeloord, The Netherlands P.O. Box 153, Introduction Tel. +31 527 248 218; Fax: +31 527 248 210 E-mail: ingen@nlr.nl; lvrijn@nlr.nl; [45, 0, -45, 0, 45, -45, 0, -45, 0, 45] spaa@nlr.nl

3 3

ABSTRACT 2.

Tensor Skin Concept

S2 [0, 45,-45, 90, 45, -45, 90, -45, dynamics and environment can be considered as a 45, 0] F2 The National Aerospace Laboratory NLR in the loop which is actively closed by the Attitude Control 4 3. Experimental ResultsNetherlands has developed a new generation of Test Computer (ACC). S3 [45, 0, -45, 45, 0, -45, 0, and Verification Equipment (TVE)0,for testing of 45, 0, 0, -45, 0, -45, 0, 45, 0, -45, 0, 45] Attitude and 4. Simulations Orbit Control Subsystems of spacecraft. 7 target attitude S4 Based on a prototype [0, 45,-45, 0, 45, 0, -45, 90, 45, 0, 0, -45, 90, -45, 0, 45, 0, -45, 45, 0] TVE developed for ESA, test measurements commands equipment has been developed for Matra Marconi CONTR L O S5 [45,-45, 0, 45,-45, 90, 0, 45, -45]s Space for AOCS subsystem performed Tests testing and system level and Future Developments 9 5. Recently of the XMM and INTEGRAL scientific-45, 0, -45, 45, -45, 45] satellites. F3 [45, -45, 45,SENSORS ACTU TORS A

3 nature of the AOCS. A complete AOCS, together with

This paper 6. Conclusions 0, concept 0, 0, -45, 45] the test F4 describes [45,-45, 0, 90, and the architecture of the XMM test system with its main features, the incremental development and delivery, References and experiences obtained during development and use of the system. The described work has also been performed under ESA II TEST MATRIX INDICATING TABLE contract. 1. INTRODUCTION 4 Tables Based on experiences with the production and S2 of use S1 8 Figures various test systems for the ISO, SAX, SOHO and F1 3 3 other satellites, the National Aerospace Laboratory NLR in the Netherlands has developed a new 3 3 generation F2 generic Test and Verification Equipment of (TVE) with re-usable hardware and software for 3 3 testing of F3 Attitude and Orbit Control Subsystems (AOCS) of spacecraft [Ref. 1].

9Info on: Earth Sun Stars Magnetic field attitude DYNAMICS disturbances ENVIR ONMENT torques magnetic fields

9

THE NUMBER OF SPECIMENS PER CONFIGURATIONFig. 1 Generic attitude contr ol system

Syst

S3

S4

S5

Sy B Int

F4

3 3 In the integration and test phase the AOCS subsystem Physical Le vel is gradually built up Attitude depending on the schedule of Data 3 3o incoming units. Verification of attitude control modes Contr l Computer AOCS and real-time behaviour is done in the early period of MA b CS us 3 3 integration using a combination of real and simulated units. Protocol Le velSensor Electr nics o Actuator Electr nics o

3

Data

The TVE had to be usable from the early stage of the AOCS development up to the integration of the AOCS in the spacecraft environment i.e. open loop tests with a single unit up to closed loop tests with any combination of real and simulated AOCS units should be supported. A prototype TVE was built for ESA/ESTEC to demonstrate the new approach with re-usable hardware and software [Ref. 2]. This prototype has recently been developed into a fullblown AOCS test system able to meet the requirements for both subsystem and system level testing of the AOCS of the XMM and INTEGRAL satellites. 2. TEST CONCEPT (9 pages in total) Figure 1 gives a schematic overview of a generic AOCS for spacecraft. The diagram reflects the cyclic

The test concept described in this paper is based on a static closed loop test facility (no real motion). The Sensor test configuration is shown in figureActuator dynamics 2. The Head User v and environment simulation is responsible for the Leel Data computation Electrical of stimuliCSfor the CS sensor Electrical and the units (Engineer ing Physical MA unit MA units) processing of uli Stimmonitor ulation from the actuator units. Stim uli Simdata Monitoring Monitoring The stimulated sensors will deliver sensor Gr measurements to the AOCS Unit Sim the MACS attitude ACC ulation via En control databus. Dynamics and vironment Simlation In the ACC the ureceived data will be T into the attitude control laws, which results in est Equipment fed commanding of the actuator units. The response of the Fig. 2 est igur measured with a monitoring device 3 A Fig. rchitectur o e actuator T confis ation units and routed back to the corresponding dynamics and environment model. In this way the loop is closed. The MACS interface has to be programmable to reflect any combination of real and simulated units. If real sensor and/or actuator units are not available they

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TEST AND VERIFICATION EQUIPMENT FOR THE ATTITUDE & ORBIT CONTROL SYSTEM OF THE XMM SATELLITE Contents by

TABLE III EXPERIMENTALSchenau, L.C.J.OF CURRENT PROGRAMME H.A. van Ingen RESULTS van Rijn, J. Spaa 3 Abstract Specimen Delamination onset moment [Nmm/mm] National Aerospace Laboratory NLR configuration Average Deviation3 1. S1F1 Introduction Tel. +31 527143.7 Fax: +31 527 248 210 149 155.8 147.5 5 248 218; E-mail: ingen@nlr.nl; lvrijn@nlr.nl; spaa@nlr.nl 143.5 S1F2 136.5 145.7 148.3 5.1 3 S1F3 147.3 147.3 152.4 2.4 ABSTRACT 2. Tensor Skin Concept nature of the AOCS. A149 complete AOCS, together with dynamics and environment can be considered as a S2F1 148.0 145.5 143.8 145.8 1.7 The National Aerospace Laboratory NLR in the loop156.3 is actively 152 by the Attitude Control which closed S2F2 Experimental Results 148.0 151.8 3.4 4 3. Netherlands has developed a new generation of Test Computer (ACC). S2F3 154.9 148.1 156.1 7.1 and Verification Equipment 165.3 for testing of (TVE) Attitude and 4. Simulations Orbit Control Subsystems of spacecraft. S3F1 673.4 588.2 515.9 592.5attitude 64.4 7 target Based on a prototype TVE developed for ESA, test S3F2 526.8 513.0 504.4 514.7 commands 9.2 measurements equipment has been developed for Matra Marconi CONTR L O S3F3 Recently 565.7 528.4 544.7 15.6 Space for AOCS subsystem performed Tests 540.0 and system level testing 9 5. and Future Developments of the XMM and INTEGRAL477.6 scientific satellites. S4F1 558.5 477.0 504.4 38.3 SENSORS ACTU TORS A S4F2 478.7 452.3 534.1 488.4 34.1 9 This paper 6. Conclusions describes the test concept and the S4F3 538.8 505.5 526.1 14.7 architecture of the XMM test system with its534.1 main DYNAMICS magnetic attitude features, S5F4 the incremental development and delivery, 531.6 631.8 576.4 torques 41.6 Info565.9 on: fields References and experiences obtained during development and use Earth Sun of the system. The described work has also been Stars performedTABLE IV LAMINATE LAY-UP DEFINITION FOR under ESA contract. Magnetic fielddisturbances

P.O. Box 153, 8300 AD Emmeloord, The Netherlands

9

SKIN AND STIFFENER ENVIR ONMENT

Syst

1. INTRODUCTION 4 Tables Fig. 1 Generic attitude contr ol system Based on experiences with the production and use of Code 8Laminate lay-up Figures various test systems for the ISO, SAX, SOHO and other 1a satellites, the 45, -45, 45, -45, 45, 45, -45, 45, -45] integration and test phase the AOCS subsystem In the [-45, National Aerospace Laboratory Physical Le vel NLR in the Netherlands has developed a new is gradually built up Attitude depending on the schedule of Data generation of generic0, 45, 0, Verification Equipment Contr l Computer o 1b [-45, Test and -45, 45, 0, -45, 0, -45] incoming units. Verification of attitude control modes AOCS (TVE) with re-usable hardware and software for and real-time behaviour is done in the early period of MA b CS us [-45, and Orbit Control Subsystems testing 2a Attitude 45, -45, 45, -45, 45, -45, 45, -45, 45]s of integration using a combination of real and simulated (AOCS) of spacecraft [Ref. 1]. units. Protocol Le vel

Skin laminates

Sy B Int

2b

[45, 0, -45, 0, 45, 0, -45, 0, 45, 0, 0, -45, 0, -45, 0, 45, 0, -45, 0, 45] SensorElectr nics o

The TVE had to be usable from the early stage of the AOCS development up to the integration of theStiffener AOCS in the spacecraft environment i.e. open loop tests with Code a single unit up to closed loopLaminate lay-up tests with any combination of real and simulated AOCS units should be supported.

The test concept described in this paper is based on a laminates static closed loop test facility (no real motion). The Sensor test configuration is shown in figureActuator dynamics 2. The Head Core laminate Flange User v and environment simulation is responsible for the Leel Data lay-up sensor units and the ing width computation Electrical of stimuliCSfor the CS (Engineer Physical MA unit MA Electrical units) [mm] processing of uli Stimmonitor ulation from the actuator units. Stim uli Simdata Monitoring Monitoring The stimulated sensors will deliver sensor Gr 1a - ACC ulation 44 A prototype [0, 45, 0, -45, 0, 90,ESA/ESTEC 45, 0] measurements to the AOCS Unit Sim the MACS attitude TVE was built for 0, -45, 0, to via En demonstrate the new approach with re-usable hardware control databus. Dynamics and vironment Simlation In the ACC the ureceived data will be 1b [0, 45, 0, This prototype has recently 0] fedTest Equipment -45, 0, 90, 0, -45, 0, 45, [05control laws, which results in ] 44.9 and software [Ref. 2]. into the attitude been developed into a fullblown AOCS test system commanding of the actuator units. The response of the [0, 45, 0, -45, 0, 90, subsystem 45, [0,-45,0,45,0] 45.81 Fig. 2 est igur measured with a smonitoring device 3 A Fig. rchitectur o e able to1c meet the requirements for both 0, -45, 0, and 0] actuator T confis ation units system level testing of the AOCS of the XMM and and routed back to the corresponding dynamics and 2a [0, 0, 45, 0, -45, 0, 90, 0, -45, 0, 45, 0, environment model. -In this way the loop50 closed. 0] INTEGRAL satellites. is 2. TEST CONCEPT (9 pages in total) Figure 3a gives a 45, -45, 45, 0, -45, of a generic 1 [0, schematic overview 0, 90, 0,-45, AOCS for spacecraft. The diagram reflects the cyclic

Actuator Electr nics o

Data

2b

[0, 0, 45, 0, -45, 0, 90, 0, -45, 0, 45, 0, 0] 0,

[0,-45,0,45,0] 51.81 The MACS interface has to be s programmable to reflect any combination 45, -45, 45, 0] - of real and simulated units. If 44 real sensor and/or actuator units are not available they 56

3b

[0, 45, -45, 45, 0, -45, 0, 90, 0,-45, 0, 45, -45, 45, 0]

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TABLE V FOUR-POINTvan Ingen Schenau, L.C.J. van Rijn, OBTAINED IN EARLIER H.A. BENDING TEST RESULTS J. Spaa 3 Abstract PROGRAMME (REF. 2) National Aerospace Laboratory NLR [Nmm/mm] Specimen Delamination onset moment P.O. Box 153, 8300 AD Emmeloord, The Netherlands configuration Deviation 3 1. Introduction Tel. +31 527 248 218; Fax: +31 527 248 210 Average 1a1a 101 E-mail: ingen@nlr.nl; lvrijn@nlr.nl; spaa@nlr.nl 105 109 105 3.3 1a1b 90.9 98.9 107 98.9 6.6 3 ABSTRACT 2. Tensor Skin Concept nature of the AOCS. A complete AOCS, together with 1a2a 104 124 112 113.3 8.2 dynamics and environment can be considered as a The National Aerospace Laboratory NLR in the loop 851 is actively100.7 by the Attitude Control which closed 1a2b Experimental Results 100 117 13 4 3. Netherlands has developed a new generation of Test Computer (ACC). 1b1a 92 874 93.1 5.2 and Verification Equipment 100 (TVE) for testing of 1b1b 92 886 886 89.7 1.6 Attitude and 4. Simulations Orbit Control Subsystems of spacecraft. 7 target attitude Based on1b2a a prototype TVE developed for ESA, test 73.6 863 886 82.8 6.6 measurements commands equipment has been developed for Matra Marconi CONTR L O 93.2 84 92.4 6.6 Space for1b2b subsystem performed Tests100 Future Developments AOCS Recently and system level testing 9 5. and of the XMM and INTEGRAL506 scientific satellites. 2a1a 454 500 486.7 23.2 SENSORS ACTU TORS A 2a1c 374 305 385 354.7 35.4 9 This paper 6. Conclusions describes the test concept and the 2a3a 472 431 475.3 37.6 architecture of the XMM test system with its523 main DYNAMICS magnetic attitude features, 2a3b the incremental development and delivery, 426 282 328 345.3 torques 60.1 Info on: fields 9 References during development and use and experiences obtained disturbances Earth 2b1a The described work has also460 437 460 452.3 10.8 Sun of the system. been Stars ENVIR ONMENT 2b1c 500 460 480 performed under ESA contract. Magnetic field 2b3a 397 466 385 416 35.7 1. INTRODUCTION 4 Tables Fig. 1 Generic attitude contr ol system 460 426 479.3 53.2 Based on2b3b experiences with the production and 552 of usevarious test systems for the ISO, SAX, SOHO and other satellites, the VI LATERAL TENSION TEST RESULTS OBTAINEDphase the AOCS subsystem In the integration and test IN EARLIER TABLE National Aerospace Laboratory Physical Le vel NLR in the Netherlands has developed a PROGRAMME new is gradually built up Attitude depending on the schedule of Data generation of generic Test and Verification Equipment incoming units. Verification of attitude control modes Contr l Computer o Specimen configuration Delamination onset real-time behaviour is done in the early period of load AOCS (TVE) with re-usable hardware and software for and MA b CS us [N/mm]integration using a combination of real and simulated testing of Attitude and Orbit Control Subsystems (AOCS) of spacecraft [Ref. 1]. 1a1a 240 units. Protocol Le vel The TVE had to be usable from the early stage of the 1a2a AOCS development up to the integration of the AOCS240 in the spacecraft environment i.e. open loop tests with220 1a2b a single unit up to closed loop tests with any 1b1a 320 combination of real and simulated AOCS units should 1b1b 340 be supported. A prototype TVE was built for ESA/ESTEC to 1b2b demonstrate the new approach with re-usable hardware280 2a1a and software [Ref. 2]. This prototype has recently480 been developed into a fullblown AOCS test system460 2a1c able to meet the requirements for both subsystem and 2a3a system level testing of the AOCS of the XMM and480 INTEGRAL satellites. 2a3b 460 2. TEST CONCEPT (9 pages in total) Figure 1 gives a2b1c schematic overview of a generic560 AOCS for spacecraft. The diagram reflects the cyclic620 2b3a

Syst

8 Figures

Sy B Int

1a1b

240

1b2a

300

The test concept described in this paper is based on a static closed loop test facility (no real motion). The Sensor test configuration is shown in figureActuator dynamics 2. The Head User v and environment simulation is responsible for the Leel Data computation Electrical of stimuliCSfor the CS