Advanced Aerospace Applications, Volume 1

Classical (FEM, BEM) structural optimisation techniques fail to solve medium high frequency dynamic problems because too many DoFs are involved and eigenvalues and eigenvectors loose the significance due to high modal density. Using a SEA model, the subsystem energies are controlled by (coupling) loss factors, under the same loading conditions. In turn, coupling loss factors (CLF) depend on physical parameters of the subsystems. The idea is to determine an approximate relation between CLF and physical parameters that can be modified in the structural optimisation process, for instance, by using Design of Experiment (DoE). Starting from this relation, an optimisation problem can be formulated in order to bring the subsystem energies under prescribed levels. A preliminary analysis of subsystem energy sensitivity to CLF can be performed to save time in looking for the approximate relationship between CLF’s and physical parameters. The approach is applied on a typical aerospace structure.

The vibro-acoustic load during launch takes a big toll on space structures. In order to simulate the dynamic loading as encountered during launch, both shaker facilities and high sound pressure reverberant rooms are used. Acoustic particle velocity sensors offer interesting new opportunities, for measuring both the applied noise field as well as the structural responses. Single particle velocity sensors in a so-called U probe can be used for very near field vibration measurements. When they are combined with a microphone in a PU probe the full sound vector can be measured. The novel perspectives of using PU probes for reverberant room testing comprise: The classical control of the noise field and the measurement of the sound pressure level, and of acoustic quantities like the reverberation time may be complemented by making reference to the total acoustic energy. Arrays of U probes can be used to measure contactless surface vibrations at multiple measurement points simultaneously as an alternative to accelerometers or laser vibrometers. PU probes can be used to measure local acoustic quantities near the structure like sound radiation, acoustic impedance and energy. 3D sound fields around structures can be visualized. The degree of diffusion in a reverberation room can be better characterized. In this paper, the first experience with practical implementation and the results of recent measurements using the PU probe will be presented.

Vibration modal analysis and ultrasonic guided wave tests are two commonly used defect detection techniques for Nondestructive Evaluation (NDE) and Structural Health Monitoring (SHM). Each technique has its unique technical merits and also limitations. In this paper, we discuss a new high frequency ultrasonic vibration test method that is developed to bridge the conventional low frequency vibration analysis together with ultrasonic guided wave techniques. The new ultrasonic vibration modal analysis technique (UMAT) takes advantage of the high defect detection sensitivity of ultrasonic guided waves, and at the same time, requires only a minimum number of testing points to cover a large structure, as can be achieved in low frequency vibration analysis. Time delay annular array actuators that are capable of tuning guided wave modes and frequencies are applied to introduce controlled high frequency vibrations to structures being tested. The defect detection sensitivity of the controlled vibrations depends on the sensitivity of the input guided wave modes and frequencies, which is associated with the displacement and stress wave structures of the guided waves. By looking at the vibration patterns and the resonant frequencies under guided wave mode and frequency tuning based on wave structure considerations, high defect detection sensitivity is achieved.

A high intensity acoustic test in a reverberant chamber was conducted on the CASSIOPE spacecraft in the final stages of integration and test campaign to ensure that it would survive the acoustic loads during launch. This paper describes the acoustic test methodology, the details of the model used for analytical prediction of the structural response for acoustic excitation and discussion of the predicted response comparison with test results that provided confidence in the spacecraft structural design for acoustic loads. The objective of the spacecraft acoustic test was to demonstrate the ability of the structure and avionics to withstand the broadband random acoustic environment experienced within the launch vehicle payload fairing. The CASSIOPE spacecraft was tested in the reverberant chamber at overall sound pressure level up to 142.1 dB. The automatic spectral control system of the acoustic test facility, which used six control microphones, was able to achieve and the maintain target spectrum levels around the spacecraft within tolerances without manual adjustments to the noise generators’ controls. The dynamic response of the CASSIOPE spacecraft during the test was measured using a large number of accelerometers installed on critical locations of the structure. Low level pre-test and post-test structural response signatures as well as electrical integrity checks performed after the exposure to the proto-flight acoustic environment demonstrated the ability of the spacecraft to survive the launch. The acoustic response of the spacecraft was also predicted based on a finite element model analysis to identify the critical components, evaluate structural margins and assess the risks in proceeding with a proto-flight acoustic test based on the specified launch vehicle spectrum. The analysis method used to predict the responses combines the NX/NASTRAN solver and RAYON, a vibro-acoustic simulation software. The RAYON software functionality is based on a boundary element model that enables the creation of an accurate fluid loading on the structure, with consideration of fluid mass and damping effects. The study used a finite element model of the structure that was correlated through an experimental modal survey test and actual spectrum levels achieved during the acoustic test. Responses of most locations compared favourably with the predictions in critical locations such as the solar arrays. Due to the limited availability of the satellite as well as time and cost constraints in a spacecraft development program, it is important to perform both qualification tests as well as analytical predictions in an efficient and timely manner to validate structural designs of spacecraft.

The Fine Guidance Sensor (FGS) is the Canadian contribution to the James Webb Space Telescope (JWST). The optical assembly (OA) is one of the components of the FGS. In October 2009, the Engineering Test Unit (ETU) of the FGS OA was tested in preparation for testing of the flight hardware planned for the Spring of 2011. The presence of a required interface ring having about 25% of the test item mass (80 kg) complicated the planning and performance of the Force Limited Vibration (FLV) testing. Modifications to our standard FLV procedure were developed and successfully applied during testing of the ETU. Such modifications were deemed necessary based on pre-test analysis showing that the FGS OA apparent or dynamic mass was significantly changed due to the presence of the interface ring. The paper first main part presents results of the pre-test analysis that demonstrated that the proposed simple modifications to the standard FLV procedure would ensure conservatism during testing (when compared to testing without the interface ring). The other main part of the paper presents details and results of the successful ETU vibration test. For one of the lateral axes, force limiting was combined with response limiting applied in the higher frequency range. The use of force limiting resulted in significant reduction of overtesting for all three axes of excitation. For the two lateral axes, the input acceleration PSD was notched by more than 20 dB; for the vertical axis, it was notched by 17 dB.

The Force Limited Vibration approach was developed in the nineties to reduce the overtesting associated with conventional vibration testing of aerospace hardware. Several methods have been considered for the estimation of the force limits. Because of its numerous advantages, the semi-empirical method is the most widely used technique for deriving these force limits. The paper first presents the mathematical relations of the semi-empirical method. The so-called C² constant is the only parameter of the method that cannot be obtained directly or from low-level preliminary runs. The paper thus continues with a detailed discussion on criteria normally used for selecting the C² constant and on the range in which it is normally expected to fall. The next section of the paper discusses the advantages of force limiting over the more traditional response limiting for notching vibration input to space hardware.

Experimental measurement of rigid body mass properties, specifically pitch inertia, of aircraft is becoming increasingly important for flutter analysis. This paper proposes a methodology for calculating rigid body inertia properties when the flexible modes of an aircraft are coupled to some degree with their corresponding rigid body modes, i.e. the rigid body modes are not fully rigid. The methodology will be demonstrated on a simple analysis model of a plate-like structure where the rigid body and primary flexible structural modes are coupled.

In most military aircraft and spacecraft applications, each payload structure must be pre-tested on a shake table to insure that it can withstand the vibration environment that it will experience during flight. Shaker testing is done using a control PSD which is designed to realistically represent the floor motion of the aircraft during takeoff, in flight, or during landing. Qualification testing is typically done by mounting the test article on one or more shakers, and exciting it with a closed loop shaker testing system so that the base of the payload responds with the pre-specified control PSD. When a test vehicle is too massive to be tested by mounting it on shakers, it is impossible to perform a base-shake test on a shake table. So the question arises; “Are there other more convenient driving points from which to excite the structure which will simulate a base-shake test?” In this approach, we derive a frequency domain Transmissibility model which is used to calculate PSDs for convenient driving points as functions of the base-shake PSDs. These calculated PSDs would then used to control a shaker test that simulates the base-shake test. The Transmissibility model is validated by using an inverse calculation to calculate base-shake PSDs as functions of thenew driving point PSDs. Suitable driving points can then be chosen by comparing the calculated base-shake PSDs with the original pre-specified base-shake PSDs.

Structural responses obtained with finite element (FE) simulations normally differ from those measured on physical prototypes. In the case of monolithic structures, the differences between the simulated and measured responses are mainly caused by inaccuracies in the geometry and material modeling. Such inaccuracies may result from the manufacturing process. The presented work illustrates how the geometry of CAD-based FE-models can be updated using a high-fidelity representation of the actual manufactured geometry, to improve the correlation between measured and computed resonant frequencies and mode shapes. The study presented in this paper was performed on a cast iron lantern housing of a gear box. In a first step, the resonant frequencies and modes shapes of the test structure were measured using impact testing. Next, a set of digital pictures were taken from a number of different angles. By means of photogrammetry, these pictures were converted into a surface model that represented the actual geometry of the lantern housing. This surface model was then compared with an FE-model derived from a CAD-model of the lantern housing. In this way, the regions where there was a substantial difference between the actual geometry and CAD-model could be identified. Finally, the geometry of the FE-model was corrected based on the measured geometry using a mesh morphing technique. For the considered test case, the correction of the geometry provided a significant improvement of the quality of FEM-test correlation of the modal parameters.

This paper presents a new approach for damage detection in structural health monitoring systems exploiting the coherence function between the signals from PZT (Lead Zirconate Titanate) transducers bonded to a host structure. The physical configuration of this new approach is similar to the configuration used in Lamb wave based methods, but the analysis and operation are different. A PZT excited by a signal with a wide frequency range acts as an actuator and others PZTs are used as sensors to receive the signal. The coherences between the signals from the PZT sensors are obtained and the standard deviation for each coherence function is computed. It is demonstrated through experimental results that the standard deviation of the coherence between the signals from the PZTs in healthy and damaged conditions is a very sensitive metric index to detect damage. Tests were carried out on an aluminum plate and the results show that the proposed methodology could be an excellent approach for structural health monitoring (SHM) applications.

Pure normal mode test, a routine ground vibration test (GVT) for large aircrafts, needs many shakers. When conducting such test for a pure mode, two major problems remain unsolved: One is how to determine the best shaker number; another is how to determine the best distribution of the shakers after shaker number is determined. This paper will answer these two important questions. When one mode is to be excited, the number of the shakers and the corresponding locations of shakers can be determined by a singular value decomposition approach, an algorithm proposed in this paper. The force appropriation can also be solved by SVD method with FRF matrix. The force appropriation effect is quantified by two indexes, the pure index and effective index. And the complete FRF matrix can be constructed by modal parameters of initial modal test or FEM. At the first section of this paper, the pure normal mode test theory development is reviewed. At the end, a simulation example is given to help understanding the new theory.

This paper presents all the updating activities performed on the finite element model of PLEIADES. The model updating is usually limited to a correction of modal data, by changing the most sensitive physical design parameters. In this paper, the modelization errors are localized and corrected thanks to a residual energy criteria: the Constitutive Relation Error (CRE). This method was originally developed by the LMT Cachan, and then implemented by the FEMTO Institute (Besançon, FRANCE) for application in an industrial context. The updating of PLEIADES is based on a modal approach: The experimental modes are identified using the Real Time Modal Vibration Identification (RTMVI) method. First, the model of the payload is updated with respect to a subsystem test performed on the instrument. Next, the model is condensed and included in the satellite model. The final step is to update the entire model using tests at satellite level. Primodal, a structural analysis tool developed by TOPMODAL (Toulouse, FRANCE) is used for correlation and updating.

Ground vibration testing (GVT), one of the critical tests which occur during aircraft development, is typically one of the last tests to take place prior to embarking on the flight test program, providing valuable information for the validation of the aeroelastic stability of the aircraft. Historically, the GVT is required by the aviation regulators in the certification process. This highly visible and time-constrained test has evolved over the years as new data collection tools, both hardware and software, have become available. The Gulfstream G650 aircraft serves as an example of how modern approaches have allowed this required test to provide highly evolved information much more efficiently and with improved confidence, dispelling the myth that testing has to be time-consuming, costly, and complicated in order to be considered a success.

The Stratospheric Observatory for Infrared Astronomy (SOFIA) is now in operation out of NASA Dryden Flight Research Center and is providing astronomical science observations not possible from other Earth- and spaceborne observatories. A 2.7 meter telescope which weighs 34,000 pounds has been installed in the aft fuselage of a Boeing 747SP aircraft. This required significant structural changes to the airframe. Further, the telescope is installed on an isolation system with a dynamic control system for properly positioning and controlling the pointing of the telescope. The major structural modifications of the aircraft and the dynamic interaction of the telescope with the aircraft are important issues related to the aircraft flight stability. Ground vibration testing (GVT) was performed to recharacterize the dynamic properties of the aircraft with the telescope installed. This testing was performed prior to delivery to NASA DFRC, where the flight test program was conducted. In addition to standard GVT measurement parameters, the testing involved evaluation of the acoustic cavity where the telescope is installed. Dynamic testing was also performed to investigate the structural coupling interaction that occurs when the telescope control system is activated.

Impact resistance of different types of composite sandwich beams is evaluated by studying vibration response changes (natural frequency and damping ratio). This experimental works will help aerospace structural engineer in assess structural integrity using classification of impact resistance of various composite sandwich beams (entangled carbon and glass fibers, honeycomb and foam cores). Low velocity impacts are done below the BVID limit in order to detect damage by vibration testing that is hardly visible on the surface. Experimental tests are done using both burst random and sine dwell testing in order to have a better confidence level on the extracted modal parameters. Results show that the entangled sandwich beams have a better resistance against impact as compared to classical core materials.

This paper presents the methodology used for the flight test conducted as part of the gross weight increase program for NASA’s WB-57F aircraft. The primary focus is on the synergy of data processing procedures to track the trends of the frequency and damping of the modes of interest in near-real time during the flight test. These procedures include spatial filtering with FEM mode shapes, time domain parameter estimation, and a pole density diagram. This method can be used to estimate modal parameters from free-decay response-only measurements using only a few acceleration response measurements on the aircraft. The authors will show that with this method, a robust estimate for frequency and damping can be extracted and statistics of the estimated values can be evaluated at each flight test point. The paper will also discuss the preliminary testing and analyses as well as the overall logistics of the flight test program, including instrumentation selection and installation, data acquisition and transfer, and operational considerations.

The identification of multiple-site damage is a challenging problem in data-based structural health monitoring (SHM). It is generally accepted that higher level damage identification via statistical pattern recognition requires the adoption of a supervised learning approach, with the need for data to be gathered from the structure in all damaged states of interest. The number of states for which data would be required to cover all damage combinations grows exponentially with the number of locations at which damage may occur. Damage state data sets of this extent are unlikely to be available in practical applications. The objective of this paper is to explore an interesting approach to the problem of multiple-site damage location. It is postulated that if sufficient information can be gleaned from single-site damage data to allow identification of multiplesite damage, then the requirement to gather data for all combinations of damage location may be circumvented. In the present study this possibility is assessed using data from an experimental structure. The experimental structure used is a full-scale, laboratory-based aircraft wing section. Damage sensitive features identified using single-site data are shown to perform well when applied to the multiple-site location problem.

In this paper, several techniques for nonlinear system identification are applied to a real-world structure, the SmallSat spacecraft structure developed by EADS-Astrium. This composite structure comprises two vibration isolation systems, one of which possesses mechanical stops. The loading case considered in the present study is a random (local) excitation. A careful progression through the different steps of the system identification process, namely detection, characterization and parameter estimation, is carried out. Different methods are applied to data resulting from numerical experiments, without having access to the finite element model which generated these data.

Ground Vibration Testing (GVT) of aircraft is performed very late in the development process. The main purpose of the test is to obtain experimental vibration data of the whole aircraft structure for validating and improving its structural dynamic models. Among other things, these models are used to predict the flutter behavior and carefully plan the safety-critical in-flight flutter tests. Due to the limited availability of the aircraft for a GVT and the fact that multiple configurations need to be tested, an extreme time pressure exists in getting the test results efficiently. The aim of the paper is to discuss modern testing and analysis concepts for performing a GVT that are able to help realize an important testing and analysis time reduction without compromising the accuracy of the results. For the past several years, LMS has organized so-called Master Classes on the GVT topic. The aim of the class is to introduce an integrated approach to handle the test preparation, modal testing, modal analysis, numerical model correlation, model updating, and model exploitation, to the industry by means of using a full-scale aircraft. This paper illustrates this approach by presenting testing and analysis results of an F-16 aircraft.

The need to reduce testing time without diminishing the quality of the data is an important driver for innovation in the aerospace testing industry. In this paper, the use of advanced, flexible shaker excitation signals will be investigated with the aim (1) to obtain improved Frequency Response Function (FRF) estimations and (2) to assess the non-linearities of the excited system / structure. Pseudo-random and more general multisine signals, rather than the more traditional pure or burst random signals, will be used to increase the accuracy of the FRF estimate. Moreover, special multisine data acquisition and processing methods to identify the level of non-linearity will be illustrated by means of Ground Vibration Testing data of an F-16 aircraft. The presented methods allow assessing the non-linearities at a single excitation level, which is in contrast to the more traditional method of repeating the test at multiple excitation levels and observing the FRF differences. In addition, a new perspective will be given on the post-processing of stepped sine FRFs. Stepped sine shaker excitation signals are traditionally used to highlight and study non-linear behaviour. In this paper, a curve-fitting method based on FRF data at fixed response levels is applied to identify and quantify the non-linearities of the structure. Again, the approach will be illustrated by means of F-16 aircraft data.

A method for compensating lateral - anti-symmetric and longitudinal - symmetric structural vibrations of an aircraft caused by turbulence, gust, wind blasts, wake penetration and buffeting in flight is presented. The method includes the steps of detecting the structural vibrations by a novel measurement technology using pitch, roll and yaw rates determined by an inertial sensing system. The noval measurement supplying the determined disturbing values to a flight control system, producing phase- and amplitude-correct control flap movements by generating appropriate control signals in respective control drives to counteract the phases and amplitudes of the excited vibrations. The design requirements especially the elastic mode stability criteria are described. For the symmetric mode alleviation additional means for improvement using high frequency trailing edge split-flap flight control system control laws are outlined. Also lateral vibration alleviation using rudder and trailing edge flaps is described.

Two mounts were added to a helicopter making it possible to carry different payloads. To validate the structural effects of these modifications, modal tests were performed on-ground on the helicopter in its standard configuration as well as in its modified configuration with the added payloads. In addition, an in-flight test was performed to verify the impact on the existing flight envelope. For all tests, Operational Modal Analysis was used. The obtained results allowed for updating the flight procedures and operating profiles for the helicopter and provided added flexibility with respect to the best possible helicopter configuration to obtain the mission objectives, while maintaining optimum safety for the flight crew.

Wavelet analysis is a powerful method for analyzing the time histories of signals. A discrete wavelet family is developed for structural dynamics by using the dilation equation to calculate scaling function coefficient values for arbitrary waveforms. The performance of this formula is verified by analyzing the scaling functions of multiple Daubechies wavelets. To assure the new discrete wavelet families have the characteristics of a specific system, the formula is applied to analytical and experimental response data. The relationship between the number of coefficients and their ability to successfully capture the characteristics of the signal is studied and a method is developed for determining the number of coefficients to be used when applying the formula. The resulting new families of discrete wavelets are based upon the nominal characteristics of a given system for use in signal processing and model discretization applications. The impulse response of a structure is proposed as a tuned baseline for structural health monitoring applications; the corresponding Wavelet Analysis of Structural Anomalies using Baseline Impulse-Response (WASABI) method is presented and discussed in the context of the wavelet development.

In dynamic analysis of structures, the accuracy of the mathematical model plays a crucial role. However, because of several uncertainties like local nonlinearities, welding points, bolted joints, material properties and geometric tolerances, the mathematical model will contain differences compared to the manufactured product. Hence, it is essential to update mathematical models by using vibration test data taken from the structure. This paper presents a new approach to model updating via utilizing neural networks and genetic optimization algorithms. The key point in this new approach is that the model updating capability of neural networks is improved by a genetic optimization algorithm by guiding the optimization problem with results obtained from neural network identification. Employing the nominal mathematical model created for a particular structure, a data set of selected mode shapes and natural frequencies is created by a number of simulations performed by perturbing selected updating parameters randomly. A neural network is then created and trained with this data set. Upon training the network, it is used to update the initial model with the test data. The results are then improved further by using the “network updated mathematical model” as an initial model and updating it again by employing a genetic optimization algorithm. The most important advantage of the proposed approach is the possibility of using different number of degrees of freedom for each mode shape; as a result, additional flexibility is introduced to the approach, since the proposed method can be used with incomplete test data. The application and capabilities of the proposed approach is illustrated via real test data taken from a GARTEUR test bed, where it is seen that the proposed method updates mathematical models associated with such complex structures efficiently.

A piezoelectric actuated stabilization mount for payloads onboard small Unmanned Aircraft Systems (UAS) is designed, analyzed, and experimentally tested. The custom three degree of freedom system is capable of producing attitude corrections in pitch and roll as well as axial stabilization. The mount assists in many applications including Intelligence Surveillance and Reconnaissance (ISR) missions, airborne target identification and tracking, and narrow beam directional communications. The system consists of a mounting plate, piezoelectric actuators, precision position sensors, and a digital controller in a tailor-made compact lightweight package. Full analytical and finite element models are created to characterize the dynamic behavior of the system. Both the modal and transient analyses are in good agreement with experimental testing results. A closed-loop Proportional, Integral, Derivative (PID) controller is implemented and tuned to actively reduce base excitation to provide a much-improved stabilized platform. The stabilization mount is subjected to single disturbances as well as random excitations over large frequency ranges for assessment of performance. Results presented include analytical, numerical, and experimental test data as well as a tuned controller for an effective piezoelectric actuated stabilization mount for small UAS.

The modeled response of spacecraft systems must be validated using flight data as ground tests cannot adequately represent the flight. Tools from the field of operational modal analysis would typically be brought to bear on such structures. However, spacecraft systems have several complicated issues: 1. High amplitudes of loads; 2. Compressive loads on the vehicle in flight; 3. Lack of generous time-synchronized flight data; 4. Changing properties during the flight; and 5. Major vehicle changes due to staging. A particularly vexing parameter to extract is modal damping. Damping estimation has become a more critical issue as new mass-driven vehicle designs seek to use the highest damping value possible. The paper will focus on recent efforts to utilize spacecraft flight data to extract system parameters, with a special interest on modal damping. This work utilizes the analysis of correlation functions derived from a sliding window technique applied to the time record. Four different case studies are reported in the sequence that drove the authors’ understanding. The insights derived from these four exercises are preliminary conclusions for the general state-of-the-art, but may be of specific utility to similar problems approached with similar tools.

The rapid deployment of satellites is hindered by the need to flight-qualify their components and the resulting mechanical assembly. Conventional methods for qualification testing of satellite components are costly and time consuming. Furthermore, full-scale vehicles must be subjected to launch loads during testing. This harsh testing environment increases the risk of component damage during qualification. The focus of this research effort was to assess the performance of Structural Health Monitoring (SHM) techniques as a replacement for traditional vibration testing. SHM techniques were applied on a small-scale structure representative of a responsive satellite. The test structure consisted of an extruded aluminum space-frame covered with aluminum shear plates, which was assembled using bolted joints. Multiple piezoelectric patches were bonded to the test structure and acted as combined actuators and sensors. Various methods of SHM were explored including impedance-based health monitoring, wave propagation, and conventional frequency response functions. Using these methods in conjunction with finite element modelling, the dynamic properties of the test structure were established and areas of potential damage were identified and localized. The adequacy of the results from each SHM method was validated by comparison to results from conventional vibration testing. LA-UR 10-07115.

A sensor is introduced that enables new methods for modal testing. Noncontact inertiareferenced velocity (NIRV) sensors may be mounted to a flexible stand without concern for measurement contamination due to motion of the stand. Non-contact sensors that are stand mounted may be setup without the test article, saving large amounts of the test article’s time, particularly for high-value articles and repetitive sequences on similar layouts. In cases where Computer Aided Design models of the test article are available, dimensions of the stand and measurement degrees-of-freedom may be identified without taking measurements directly from the test article. Specifications of the sensor are presented along with results comparing its output to that of traditional accelerometer data and examining differences in reduction methods. Data is presented that demonstrates insensitivity to motion at the NIRV sensor mount. Ground Vibration Test (GVT) results are presented from an aircraft with both NIRV and accelerometer data to demonstrate the interchangeability.

Launcher components, large payloads – and particularly in the past – also all major Space Shuttle payloads have required modal survey tests to determine the modal parameters with high accuracy and reliability to validate the respective dynamic structural models for coupled loads analysis and flight piloting, but also for system level structural qualification. The modal survey test methods applied usually consist of the classical ground vibration test using tuned sine excitation and the modal analysis of measured frequency response functions. While the tuned sine excitation directly results in well understood and established modal parameters – typically of the fundamental or primary modes, the modal analysis is used to complement the modal model with respect to the secondary modes. This two-folded approach serves both for high accuracy of the modal model considering the fundamental modes as well as to minimize test effort to complement the modal model in the higher frequency bands. Almost all available modal analysis tools – besides of very specific ones – which are available today are based on linear models and on a more or less statistical approach to estimate the modal parameters. With respect to real measured transfer functions, in particular when modal density increases and some non-linearity is present, which is typical of large space structures, the modal analysis results inherently suffer from a considerable amount of scatter or uncertainty. In this paper, application cases are studied to provide insight in the actual challenges to extract reliable modal parameters from large spaceflight components. Based on this, a consolidated strategy is discussed to extract the best match between experimental data and experimental modal model. In addition, the remaining uncertainty may be quantified as well.

CASSIOPE is a small Canadian spacecraft carrying two payloads: the e-POP payload consisting of eight scientific instruments and the CASCADE payload - a high speed Ka-Band communication technology demonstrator. As part of the Cassiope environmental test campaign, the spacecraft was subjected to sine vibration testing for structural qualification, and to random vibration testing for workmanship verification. The purpose of the spacecraft sine vibration test was twofold: subject the spacecraft to protoflight launch environment and to provide a test verified structural model for final coupled loads analysis with the launch vehicle. In this paper, we employ an operational modal analysis technique to the ground vibration sine qualification test of the Cassiope spacecraft. The purpose of this exercise was to determine the applicability of operational modal processing to vibration qualification tests in cases where interface forces between the vibration shaker and the test article are not able to be determined in the course of the test and traditional modal techniques are not applicable. This paper will examine the ability of operational modal techniques in separating complex vibration dynamics and modal interactions present in sine vibration qualification tests and in aiding the determination of required test notching. Cassiope (CASCADE Smallsat and Ionospheric Polar Explorer) is a Canadian smallsat scheduled to be launched in 2011[1]. The three goals of the Casssiope mission are to demonstrate the high speed Ka-Band store and forward capability of the CASCADE CX payload (provided by MacDonald, Dettwiler and Associates Ltd.) [2], to investigate the atmospheric and plasma flows and related wave particle interaction and radio wave propagation in the topside ionosphere with a suite of eight e-POP (Enhanced Polar Outflow Probe) science instruments provided by the University of Calgary, and to develop a Canadian Smallsat Bus (provided by Bristol Aerospace Ltd). Development programs for space hardware generally envision two phases of testing. The first stage would be a modal survey test consisting of excitation by multiple small shakers and optimized response channels throughout the spacecraft. The purpose of this test is to allow updating of the finite element models to allow for better estimates of the flight loads through coupled loads analysis and also better predictions of the vibration test response levels and estimation of notching profiles. However, in many cases, due to budgetary or schedule constraints, the modal test is often omitted. The second stage of testing is qualification testing consisting of sine vibration testing, random vibration testing and acoustic testing. The sine vibration testing is closed loop controlled over a limited frequency band, with notching appropriate for the loads predicted during launch (these may be base shear or bending moment, c.g. acceleration, etc). The random vibration testing is often called for as a workmanship test or to capture low frequency content not easily excited by an acoustic test. This paper first looks at an overview of the Cassiope spacecraft. This is followed by an outline of the vibration test setup and sequence. The next section looks at the use of operational modal technique in processing the sine vibration test data. The modes in each axis are determined and their interactions are discussed. The applicability of the modal results to aid in the prediction of sine test notching is presented.