AN OVERVIEW OF THE E.C.S.S. HANDBOOK FOR SPACECRAFT LOADS ANALYSIS

Transcription

1 COMPDYN 2011 III ECCOMAS Thematic Conference on Computational Methods in Structural Dynamics and Earthquake Engineering M. Papadrakakis, M. Fragiadakis, V. Plevris (eds.) Corfu, Greece, May 2011 AN OVERVIEW OF THE E.C.S.S. HANDBOOK FOR SPACECRAFT LOADS ANALYSIS Adriano Calvi 1 1 European Space Agency Keplerlaan 1, Noordwijk, The Netherlands Adriano.Calvi@esa.int Keywords: Loads Analysis, Spacecraft Structure, Structural Dynamics, Standardization, ECSS, Handbook. Abstract. The handbook of the European Cooperation for Space Standardization (ECSS) E- HB titled Guidelines for Loads Analysis of Spacecraft and Payloads is developed with the goal to harmonize methodologies, procedures and practices currently applied for the conduct of spacecraft and payloads loads analysis. The handbook will make available to the European Space Community a set of well proved methods and procedures for the calculation and assessment of structural design loads and for the evaluation of the test loads. The handbook is presently drafted by the Loads Analysis Handbook Working Group and contains a number of topics strictly related to computational structural dynamics. An overview of the proposed handbook, with a synopsis of some chapters, has been developed and it is here presented. It is planned that the handbook be published in the year Comments on the proposed handbook are solicited from the scientific community and industry.

2 1 INTRODUCTION European Cooperation for Space Standardization (ECSS) [1] is an initiative established to develop a single set of consistent space documents for use by the entire European space community. The objective is to increase the effectiveness of all space programs in Europe through the application of the ECSS documents recognized by all potential European customers and equally accepted by industry. The ECSS documents have evolved from the previous 30 years of experience in the management and implementation of European space projects. ECSS is supported by several agencies and companies. The actual members of ECSS are the European Space Agency (ESA), the National Space Agencies of Canada, France, Germany, Italy, Norway, the Netherlands, the United Kingdom, and the European space industry represented by Eurospace. The ECSS handbooks are nonnormative documents providing background information, orientation, advice or recommendations. A handbook therefore contains advice about how to properly do something important and useful information about a subject. The ECSS E-HB Loads Analysis Handbook [2] is developed with the goal to harmonize methodologies, procedures and practices currently applied for the conduct of spacecraft and payloads loads analysis. It will make available to the European Space Community a set of well proved methods and procedures for the calculation and assessment of structural design loads and for the evaluation of the test loads. The handbook is presently drafted by the Loads Analysis Handbook Working Group and contains a number of topics related to computational structural dynamics. In particular, recent advances in the area of structural dynamics and vibrations, in both methodology and capability, have the potential to make spacecraft system analysis and testing more effective from technical, cost, and hardware safety points of view. However, application of advanced analysis methods varies among the Space Agencies and their contractors. Identification and refinement of the best of these methodologies and implementation approaches is an objective the Working Group. An overview of the proposed handbook, with a synopsis of some chapters, has been developed and it is here presented. 2 GENERAL CONTENT OF THE HANDBOOK The handbook is organized in a number of chapters which cover the following topics: 1. Terms and definitions 2. General aspects of the loads verification 3. Loads combination 4. Launcher/spacecraft coupled loads analysis (CLA) 5. Static loads 6. Sine vibration environment 7. Random vibration and vibro-acoustic environments 8. Shock environment 9. Structural stability 10. Dimensional stability 11. Fatigue and fracture control 12. Microgravity and micro-vibration environment 13. Mathematical models Some topics are briefly illustrated in the following sections. 2

3 3 SPACECRAFT FLIGHT ENVIRONMENTS AND DYNAMIC LOADS There are three basic types of flight environments that generate dynamic loads on payload components [7]: a) The low-frequency dynamic response, typically from 0 to 100 Hz, of the launch vehicle/payload system to transient flight events. b) The high-frequency random vibration environment, which typically has significant energy in the frequency range from 20 Hz to 2000 Hz, transmitted from the launch vehicle to the payload at the launch vehicle/payload interfaces. c) The high frequency acoustic pressure environment, typically from 20 Hz to 8000 Hz, inside the payload compartment. The payload compartment acoustic pressure environment generates dynamic loads on components in two ways: (1) by direct impingement on the surfaces of exposed components, and (2) by the acoustic pressure impingement upon the component mounting structures, which induces random vibrations that are mechanically transmitted to the components. Combinations of these loads may occur at different times in flight. 4 SPACECRAFT-LAUNCHER COUPLED LOADS ANALYSIS The structural response of the spacecraft to transient flight events (low frequency mechanical environment) is simulated by spacecraft-launcher coupled dynamic analysis [3]. This is normally a transient analysis performed by using the finite element models (FEM) of the satellite and launcher, merged together, and by applying the forcing functions for the different launch events (figure 1). The main objective of the CLA is to calculate the loads on the spacecraft, where the term loads refers to the set of internal forces, displacements and accelerations that characterise the structural response to the applied forces. The loads of the spacecraft derived from the analysis are taken as a basis to verify the dimensioning of the spacecraft itself. Spacecraft programs typically perform a number of analysis loops (called loads cycles ), for example one each for preliminary design, final design, and final verification. The latter is done with test verified mathematical models. The low frequency domain typically ranges from 0 to up 100 Hz and corresponds to the frequency content of the forcing functions used in the CLA. The excitation may be of aerodynamic origin (wind, gust, buffeting at transonic velocity) or may be induced by the propulsion system (thrust build up or tail-off transient, acoustic loads in the combustion chambers, etc.). Of primary interest are the spacecraft interface accelerations and interface forces. The interface accelerations can be used to derive an equivalent sine spectrum at the spacecraft interface. The interface forces can be employed to calculate the equivalent accelerations at the spacecraft centre of gravity. Of large interest is also the recovery of the internal responses which are used to verify the structural integrity of the spacecraft and its components. The computed responses and their deduced minimum and maximum levels can be employed within the design, verification and test phases of the spacecraft. For example, secondary structures and flexible components such as solar arrays, booms, instruments and propellant tanks must also be designed (and test verified) to withstand the dynamic environment induced at the base of the spacecraft. The dynamic loads (accelerations, forces, stresses, etc.) on these components can be verified directly by means of the CLA (apart from acoustic loads under the fairing which are analysed separately). In the test verification phase of the spacecraft, the equivalent sine spectrum computed by means of CLA is used to locally reduce the prescribed spectrum from the launcher user s 3

4 manual at specific resonant frequencies. This might be required to avoid possible damage to the spacecraft structure itself or its components (solar arrays, booms, etc.). The chapter on launcher/spacecraft coupled loads analysis includes: Description of load cases Methodology Static and dynamic contributions Generation of reduced mathematical models and output transformation matrices Damping modelling Uncertainty factors Output and results evaluation Sensitivity analysis Non-linearity Figure 1: GOCE satellite FEM merged with VEGA launcher FEM for CLA (source [6]). 5 SINE VIBRATION ENVIRONMENT In a number of situations sine-wave excitations are used for qualification and protoflight testing of space vehicle hardware, even though the mission dynamic excitation being simulated is not periodic [4]. In particular the low frequency transient is often simulated at the subsystem and system assembly level using a swept-sine vibration test over a frequency range up to about 100 Hz. The magnitude and sweep rate for the resulting vibration are selected supposedly to cause the hardware response to be similar to the response predicted for the transient. Some common procedures to derive such a test are reported in [2, 4]. The use of a swept-sine excitation to simulate a transient excitation can result in the unique situation of causing a simultaneous undertest and overtest of the hardware. The undertest is due to exciting only one hardware resonance at a time during the sweep-sine test, as opposed to the simultaneous excitation of multiple resonances of the hardware, as would be induced by 4

5 the transient excitation. The potential overtest is due to applying a larger number of stress cycles to the hardware during the swept-sine test than occurs during the transient excitation. Of course, the amount of overtesting can be reduced by increasing the sweep rate. However the main cause of potential overtesting is the difference in structure boundary conditions between test and flight configurations. During a vibration test, the test article is usually attached to a very rigid fixture and it is excited or driven along a single linear direction, with the structure being completely restrained along the other five degrees-of-freedom (DOF). This generates a response of the spacecraft in its clamped natural modes and not in the coupled launcher-spacecraft modes as exhibited during the flight [5, 6]. In the flight configuration the satellite is attached to a mounting structure (i.e. adapter and launcher) that normally exhibits some flexibilities in all six DOF in the frequency range of interest. The flexibility difference in the direction of excitation is the main contributor to the overtesting phenomenon. In the flight configuration, the acceleration at the interface between the mounting structure and the test article drops at certain frequencies, resulting in valleys in the acceleration spectra. These frequencies correspond to the resonance frequencies of the test article when attached to a rigid support (such as a shaker). This phenomenon is known as the vibration absorber effect. In other words, during a vibration test, the structure is excited with a specified input acceleration that is the envelope of the flight interface acceleration, despite the amplitude at certain frequencies drops in the flight configuration. This results in exaggerated amplification of input forces, and internal stresses, at the resonance frequencies of the test article. In practice, because the input spectrum does not consider the actual flight behaviour and excites the spacecraft modes, the structural response of the spacecraft would exceed the spacecraft design capability and the values encountered during the launch simulations. To avoid such effects, the sine test input spectrum needs to be notched. However, it should be clear that the criteria used for notching do not have to generate an input spectrum that can jeopardize the flightworthiness of the spacecraft. For example this can happen if the testing levels are not coherent with the levels required by the CLA. As a general rule, the notching should be kept to a minimum and the input spectrum to be applied to the spacecraft as much as possible similar to the one requested by the launcher authority. The reason is that this spectrum provides robustness to the spacecraft design compliant with it since it covers a wide range of responses found in past spacecraft-launcher CLA and in addition it provides a sweep over the frequency band of interest, therefore covering possible deviations of local modes. It should be noted that the dynamic environments for space vehicle hardware are typically multiple-axis, i.e., the excitations occur simultaneously along all three orthogonal axes of the hardware. Acoustic tests naturally simulate a multiple-axis excitation, but shock and vibration test facilities are commonly uniaxial (figure 2). Multiple axis test facilities designed to simulate low frequency shock and vibration environments (generally below 100 Hz) are available (figure 3). For space vehicle hardware, however, it is more common to perform shock and vibration tests using machines that apply the excitation sequentially along one axis at a time. The potential error caused by simulating a multiple-axis shock and/or vibration excitation with sequentially applied single-axis excitations is widely debated. The chapter on sine vibration environment includes: Source of the sine environment Response analysis Environment and test specifications for instruments and components Methods and procedures for notching of the input spectrum in sine vibration test 5

6 Figure 2: GOCE satellite on large slip table at ESTEC (courtesy of ESA). Figure 3: Herschel satellite on Hydra at ESTEC (courtesy of ESA). 6 RANDOM VIBRATION AND VIBRO-ACOUSTIC ENVIRONMENTS Some load environments must be treated as random phenomena, when the forces involved are controlled by non-deterministic parameters [7]. Examples include high frequency engine thrust oscillation, aerodynamic buffeting of fairing, and sound pressure on the surfaces of the payload. Random vibration analysis describes the forcing functions and the corresponding structural response statistically. It is generally assumed the phasing of vibration at different frequencies is statistically uncorrelated. The amplitude of motion at each frequency is described by a 6

7 power spectral density function. In contrast to transient analysis which predicts time histories of response quantities, random vibration analysis generates the power spectral densities of these response quantities. From the power spectral density, the root mean square (rms) amplitude of the response quantity is calculated. The root-mean-square (rms) acceleration is the square root of the integral of the acceleration PSD over frequency. Random vibration limit loads are typically taken as the 3- sigma load (obtained by multiplying the rms load by 3). The most appropriate measure of the severity of a random vibration environment is the maximum PSD value or the PSD value at the frequency of the resonances of the structural item. It is a common mistake to use the rms value of the input as a measure of its severity. The problem with the rms value is that it depends strongly on the values of the PSD at very high frequencies and on the upper frequency limit, which are often irrelevant. The chapter on random vibration and vibro-acoustic environments includes: Source of the environment Response analysis Environment and test specifications for instruments and components Notching of the input spectrum in random vibration test Methodology and procedures for vibro-acoustic analysis 7 MICRO-GRAVITY AND MICRO-VIBRATION ENVIRONMENT One of the major goals for the future space utilization projects is conducting activities, experiments and processes in a very low gravity environment, the so-called micro-gravity environment [2]. This micro-gravity environment is usually made available in low-earth-orbiting spacecraft systems and pressurized module compartments over long period of days and months without interruptions. Consequently the spacecraft systems should be designed and should be operated such that limit acceleration levels are not exceeded during the performance of experiments and processes. The chapter on microgravity and micro-vibration environment includes: General aspects related to the micro-gravity environment Identification of the microgravity disturbance sources Derivation of microgravity specifications Verification of the microgravity requirements 8 MATHEMATICAL MODELS This chapter addresses the mathematical models of space structures with a special emphasis on the finite element models used for loads analysis. In particular it provides some guidelines for ensuring finite element analysis quality, i.e. the correct use of this specific technology, the finite element method - and the acceptance of the results. The handbook promotes the verification and validation (V&V) guidance proposed in the ASME Guide for V&V [8]. The chapter includes the following topics: Requirements for structure mathematical models (e.g. [9, 10]) Introduction to verification and validation in computational mechanics [8] Uncertainty quantification during design and verification loads cycles Model verification and quality assurance for spacecraft finite element analysis Mathematical model validation 7

8 9 CONCLUSIONS Recent advances in the area of structural dynamics and vibrations, in both methodology and capability, have the potential to make spacecraft system analysis and testing more effective from technical, cost, and hardware safety points of view. However, application of advanced analysis methods varies among the Space Agencies and their contractors. The ECSS Handbook for spacecraft loads analysis [2] is presently in the drafting phase and is developed with the goal to harmonize methodologies, procedures and practices currently applied for the conduct of spacecraft and payloads loads analysis. The handbook will make available to the European Space Community a set of well proved methods and procedures for the calculation and assessment of structural design loads and for the evaluation of the test loads. It is planned that the handbook be published in the year Comments on the proposed handbook are solicited from the scientific community and industry. Acknowledgments The work described in this paper was carried out by the members of the E-HB Loads Analysis Handbook Working Group. Their contributions are gratefully acknowledged. The handbook is funded by the European Cooperation for Space Standardization. REFERENCES [1] European Cooperation for Space Standardization, [2] Guidelines for loads analysis of spacecraft and payloads, ECSS Handbook E-HB-32-26, Draft Issue (To Be Published), Noordwijk, The Netherlands [3] S. Fransen, Methodologies for launcher-payload coupled dynamic analysis PhD Thesis, Technical University of Delft, The Netherlands, 2005, ISBN [4] Dynamic Environmental Criteria, NASA-HDBK-7005, March 2001 [5] Y. Soucy, A. Coté Reduction of Overtesting during Vibration Tests of Space Hardware Canadian Aeronautics and Space Journal, Vol. 48, No 1, March 2002 [6] Calvi A., Nali P. Some Remarks on the Reduction of Overtesting during Base-Drive Sine Vibration Tests of Spacecraft Proc. of the ECCOMAS Conf. on Computational Methods in Structural Dynamics and Earthquake Engineering - Rethymno, Crete, Greece, June 2007 [7] Load Analyses of Spacecraft and Payloads, NASA-STD-5002, June 1996 [8] Guide for Verification and Validation in Computational Solid Mechanics, ASME V&V , New York, NY, USA, 2006 [9] Space engineering, Structural finite element models, ECSS-E-ST-32-03C, Noordwijk, The Netherlands, July 2008 [10] Space engineering, Modal survey assessment, ECSS-E-ST-32-11C, Noordwijk, The Netherlands, July