Thesis (MEng) Stellenbosch University, 2014 / ENGLISH ABSTRACT: An Attitude Determination and Control System (ADCS) demonstrator and testing platform was required for satellite engineering students. The ADCS demonstrator and testing platform will allow students to develop insight into the concepts and challenges of ADCS design and implementation. The existing model nano-satellite EyasSAT was used as a design platform for a new ADCS demonstrator. A new ADCS module (ADCS_V2) was developed to replace the existing EyasSAT ADCS module. The new module allows for three-axis ADCS and the demonstration of the ADCS on an air bearing platform. The air bearing allows full freedom of movement for yaw rotations with limited pitch and roll rotations. The actuators and sensors required for the ADCS were developed and integrated into EyasSAT. In addition a new PCB was designed to form the ADCS_V2 module. Attitude determination algorithms and attitude control algorithms were implemented and tested using MATLAB Simulink simulations. These algorithms were then implemented on the ADCS_V2 module. The ADCS was tested using Hardware In the Loop (HIL) techniques and an air bearing. The yaw attitude of EyasSAT could be controlled within 0.4 degrees accuracy with all the sensors active. In order to stabilize the air bearing platform, the pitch and roll angles were rate controlled. The pitch and roll rates were damped to within 6 mrad/s. / AFRIKAANSE OPSOMMING: ’n Oriëntasiebepaling en Beheerstelsel (OBBS) demonstrasie en toets platform was benodig vir satellietingenieurswese studente. Die nuwe OBBS sal studente toelaat om insig te ontwikkel met betreking tot die idees en uitdagings wat verband hou met die ontwikkeling en implementering van ’n OBBS. Die huidige nano-sateliet model EyasSAT was gebruik as ’n ontwerpsbasis vir die nuwe OBBS. Die nuwe OBBS was ontwikkel om die huidige module van EyasSAT te vervang. Die nuwe OBBS laat oriëntasiebepaling en -beheer in drie asse toe. Die nuwe OBBS en EyasSAT kan die werking van ’n OBBS demonstreer op ’n luglaerplatform. Die luglaer laat vrye rotasie om die gierhoek toe terwyl die rol- en stygings-as beperk word. Die aktueerders en sensors wat benodig word vir die OBBS is ontwikkel en geïntegreer in EyasSAT saam met ’n nuwe gedrukte stroombaanbord om die nuwe OBBS te vorm. Orientasiebepaling en orientasiebeheer algoritmes is geïmplementeer en getoets met die hulp van MATLAB Simulink simulasies. Die algoritmes was op die OBBS module geïmplementeer en getoets deur gebruik te maak van HIL tegnieke en praktiese toetse op die luglaer. Die rotasie hoek van EyasSAT kan met ’n akkuraatheid van 0.4 grade beheer word indien al die sensors gebruik word. Die rol en stygingshoeksnelheid was gekanselleer om die luglaer stabiel te hou. Die hoeksnelheid van die twee asse kon tot kleiner as 6 mrad/s beheer word.

In this thesis, a simulation harness is constructed in Simulink for the purpose of validating the Attitude Determination and Control System (ADCS) on the APTAS student CubeSat in support of the upcoming flight readiness review. The simulation results are used to verify the compliance of a subset of the requirements for the ADCS, detailed in table 1. Calibration of the onboard sun sensor array, which is used to find the sun vector for the attitude determination, is performed using a break-out board of sun sensors tested in a sun simulator. The data gathered from this test is used to model the sun sensor system in the simulation and in flight software. The results show that the sun sensor system is able to find the sun vector with an average error angle of 5.4 degrees, though the error angle may spike up to 18 degrees in operation. It is found that the complete ADCS is able to guide the spacecraft toward the desired nadir-facing attitude, though not with the accuracy specified in the requirements. The spacecraft is able to detumble much better than required. All deficiencies found in the ADCS software have been corrected. These changes arelisted in appendix B. It is concluded that, despite its flaws, the ADCS software is flight ready. / Project APTAS

This dissertation addresses novel techniques in determining gyroless micro-satellite attitude and attitude rate. The main objective of this thesis is to explore the possibility of using commercially available low cost micro-light star sensor as a stand-alone sensor for micro-satellite attitude as well as attitude rate determination. The objective is achieved by developing accurate and computationally efficient algorithms for the realization of onboard operation of a low fidelity star sensor. All the algorithms developed here are tested with the measurement noise presented in the catalog of the sensor array STAR-1000. A novel accurate second order sliding mode observer (SOSMO) is designed for discrete time uncertain linear multi-output system. Our design procedure is effective for both matched and unmatched bounded uncertain ties and/or disturbances. The bound on uncertainties and/or disturbances is assumed to be unknown. This problem is addressed in this work using the second order multiple sliding modes approach. Second order sliding manifold and corresponding sliding condition for discrete time system is defined similar on the lines of continuous counterpart. Our design is not restricted to a particular class of uncertain (matched) discrete time system. Moreover, it can handle multiple outputs unlike single out-put systems. The observer design is achieved by driving the state observation error and its first order finite difference to the vicinity of the equilibrium point (0,0) in a finite steps and maintaining them in the neighborhood thereafter. The estimation synthesis is based on Quasi Sliding Mode (QSM) design. The problem of designing sliding mode observer for a linear system subjected to unknown inputs requires observer matching condition. This condition is needed to ensure that the state estimation error is a asymptotically stable and is independent of the unknown input during the sliding motion. In the absence of a matching condition, asymptotic stability of the reduced order error dynamics on the sliding surface is not guaranteed. However, unknown bounded inputs guarantee bounded error on state estimation. The QSM design guarantees an ultimate error bound by incorporating Boundary Layer (BL) in its design procedure. The observer achieves one order of magnitude improvement in estimation accuracy than the conventional sliding mode observer (SMO) design for an unknown input. The observer estimation errors, satisfying the given stability conditions, converge to an ultimate finite bound (with in the specified BL) of O(T2), where T Is the sampling period. A relation between sliding mode gain and boundary layer is established for the existence of second order discrete sliding motion. The robustness of the proposed observer with respect to measurement noise is also analyzed. The design algorithm is very simple to apply and is implemented for two examples with different classes of disturbances (matched and unmatched) to show the effectiveness of the design. Simulation results show the robustness with respect to the measurement noise for SOSMO. Second order sliding mode observer gain can be calculated off-line and the same gain can work for large band of disturbance as long as the disturbance acting on the continuous time system is bounded and smooth. The SOSMO is simpler to implement on board compared to the other traditional nonlinear filters like Pseudo-Linear-Kalman-filter(PLKF); Extended Kalman Filter(EKF). Moreover, SMO possesses an automatic adaptation property same as optimal state estimator(like Kalman filter) with respect to the intensity of the measurement noise. The SMO rejects the noisy measurements automatically, in response to the increased noise intensity. The dynamic performance of the observer on the sliding surface can be altered and no knowledge of noise statistics is required. It is shown that the SOSMO performs more accurately than the PLKF in application to micro-satellite angular rate estimation since PLKF is not an optimal filter. A new method for estimation of satellite angular rates through derivative approach is proposed. The method is based on optic flow of star image patterns formed on a star sensor. The satellite angular rates are derived directly from the 2D-coordinates of star images. Our algorithm is computationally efficient and requires less memory allocation compared to the existing vector derivative approaches, where there is also no need for star identification. The angular rates are computed using least square solution method, based on the measurement equation obtained by optic flow of star images. These estimates are then fed into discrete time second order sliding mode observer (SOSMO). The performance of angular rate estimation by SOSMO is compared with the discrete time First order SMO and PLKF. The SOSMO gives the best estimates as compared to the other two schemes in estimating micro-satellite angular rates in all three axes. The improvement in accuracy is one order of magnitude (around1.7984 x 10−5 rad/ sec,8.9987 x 10−6 rad/ sec and1.4222 x 10−5 rad/ sec in three body axes respectively) in terms of standard deviation in steady state estimation error. A new method and algorithm is presented to determine star camera parameters along with satellite attitude with high precision even if these parameters change during long on-orbit operation. Star camera parameters and attitude need to be determined independent of each other as they both can change. An efficient, closed form solution method is developed to estimate star camera parameters (like focal length, principal point offset), lens distortions (like radial distortion) and attitude. The method is based on a two step procedure. In the first step, all parameters (except lens distortion) are estimated using a distortion free camera model. In the second step, lens distortion coefficient is estimated by linear least squares (LS) method. Here the derived camera parameters in first step are used in the camera model that incorporates distortion. However, this method requires identification of observed stars with the catalogue stars. But, on-orbit star identification is difficult as it utilizes the values of camera calibrating parameters that can change in orbit(detector and optical element alignment get change in orbit due to solar pressure or sudden temperature change) from the ground calibrated value. This difficulty is overcome by employing a camera self-calibration technique which only requires four observed stars in three consecutive image frames. Star camera parameters along with lens (radial and decentering) distortion coefficients are determined by camera self calibration technique. Finally Kalman filter is used to refine the estimated data obtained from the LS based method to improve the level of accuracy. We consider the true values of camera parameters as (u0,v0) = (512.75,511.25) pixel, f = 50.5mm; The ground calibrated values of those parameters are (u0,v0) =( 512,512) pixel, f = 50mm; Worst case radial distortion coefficient affecting the star camera lens is considered to be k1 =5 x 10−3 .Our proposed method of attitude determination achieves accuracy of the order of magnitude around 6.2288 x 10−5 rad,3.3712 x 10−5 radand5.8205 x 10−5 rad in attitude angles φ,θ and ψ. Attitude estimation by existing methods in the literature diverges from the true value since they utilize the ground calibrated values of camera parameters instead of true values. To summarize, we developed a formal theory of discrete time Second Order Sliding Mode Observer for uncertain multi-output system. Our methods achieve the desired accuracy while estimating satellite attitude and attitude rate using low fidelity star sensor data. Our methods require lower on-board processing requirement and less memory allocation; thus are suitable for micro-satellite applications. Thus, the objective of using low fidelity star sensor as stand-alone sensor in micro-satellite application is achieved.