Erwin Mooij

Tether Dynamics Analysis and Guidance and Control Design for Active Space-Debris Removal

Recent years have seen a steep increase in research being performed toward active space-debris removal. In particular, the ESA has proposed a mission in which a robotic chaser satellite would use a tether to interface with the derelict Envisat and deorbit it. This study focuses on the preliminary design of a guidance and control system to achieve this, as well as on the influence of different tether parameters on mission performance. The lumped-mass model was used to model the tether, and the influence of the number of nodes used was investigated. Then, the mission performances of nine combinations of tether length, stiffness, and damping were evaluated. This was done using a sliding-mode controller for closed-loop relative orbit control and attitude control of the chaser satellite, the performance of which was compared against a linear-quadratic regulator. An open-loop throttle-control system was used for the main engines, for which three different thrust levels were considered. It was found that higher ...

Adaptive Disturbance-Based High-Order Sliding-Mode Control for Hypersonic-Entry Vehicles

In this paper, an adaptive, disturbance-based sliding-mode controller for hypersonic-entry vehicles is proposed. The scheme is based on high-order sliding-mode theory, and is coupled to an extended...

SHEFEX-3 Optimal Feedback Entry Guidance

SHEFEX is a DLR-led series of missions for scienti c experiments and reentry technology development. SHEFEX-2 was successfully launched from Norway (Andoya Rocket Range) in June 2012. To go on with the e ort to increase the technological level for real space missions, a new challenge in the next years with the development of SHEFEX-3 arises. SHEFEX-3, foreseen to be launched in 2016, will be more complex than SHEFEX-2 in virtue of the presence of a real guided re-entry phase, while for SHEFEX-2 an autonomous Guidance and Control phase was only partially foreseen. As a consequence, the mission will be ambitious, especially in the development of the GNC subsystem. DLR GNC Systems Department will be responsible for the development of Guidance and Navigation modules, while Control will be developed by Airbus Defense and Space, in cooperation with DLR. In this work the development of the nominal entry guidance, based on the use of PseudoSpectral Methods, is discussed. This feedforward control is then coupled with a Gain-Scheduled LQR tracking controller to reduce the error on the terminal points of the mission. Results show that the proposed approach meets the requirements on the physical constraints and the terminal states, satisfying at the same time the strong limitations coming from the need to have a highly-constrained angle of attack pro le.

Optimal Terminal-Area Strategies and Energy-Tube Concept for a Winged Re-Entry Vehicle

A guidance algorithm for the Terminal Area of a winged re-entry vehicle can use reference trajectories or predicting capabilities. In both cases, it is important to know the limits of the vehicle capabilities and the best strategy for a ight in the vertical plane. This paper describes optimal trajectories and strategies for both a maximum-range and a maximum-dive ight. O-nominal

Adaptive Heat-Flux Tracking for Re-entry Guidance

To limit the mass of the vehicle’s thermal protection system, an optimal trajectory that minimizes the total integrated heat load should be flown. In essence this boils down to tracking the maximum heat-flux constraint for as long as possible. A straight-forward and simple implementation for this tracking system could be a (linear) output-feedback controller that has a fast response, although its robustness could be doubtful due to insufficient damping. A possible good alternative is a guidance-tracking system based on so-called simple adaptive control. Such a system can have excellent performance under the influence of rather large uncertainties, although its transient response can be sluggish at times. In this paper the performance of both tracking systems has been compared, as well as an integrated implementation to see whether the individual strong points of the tracking systems can be combined. The system under consideration is a hypersonic test vehicle that has to track a stagnation heat-flux constraint of 1,700 kW/m. The results show that for a nominal mission the performance of the two individual systems is equal. A Monte-Carlo analysis indicates that the tracking error is smaller for the output-feedback controller, but due to its longer tracking time the total heat load is smaller for the adaptive system. Integrating the two systems yields a significant reduction of the tracking error, albeit at the expense of a larger guidance effort.

Guidance for Autonomous Precision Landing on Atmosphereless Bodies

Two key technologies that will likely fly on next-generation planetary landers are hazard detection and avoidance (HDA) and vision-based navigation. The purpose of HDA is to make previously unsafe landing sites accessible by future vehicles, and to increase the safe landing probability overall. Vision-based navigation will enable to target specific landing sites more precisely and also plays an important role for HDA. This paper formulates requirements on the guidance system that derive from the needs of these new systems. After the introduction of a reference mission scenario with HDA and vision-based navigation inthe-loop, an extensive literature survey of the state-of-the-art in approach-phase guidancealgorithms is presented. This presents the methods under the angle of HDA-compatibility. The sheer number of papers mentioned here leads to the conclusion that trade-offs are challenging. It is recommended to do a down-selection from the papers presented here, and do a performance-based trade study for the particular test case under consideration with a few candidates. As an example, E-Guidance is presented for a Mercuryand a Moon-landing scenario. Because the results differ dramatically, it is concluded that it is essential to perform trade-offs on the actual mission scenario under study, and that trades cannot merely be based on references in the literature. Because the proposed E-Guidance method is not satisfyingly robust, it is a goal for future research to improve upon this.

Continuous Aerodynamic Modelling of Entry Shapes

During the conceptual design phase of a re-entry vehicle, the vehicle shape can be varied and its impact on performance evaluated. To this end, the continuous modeling of the aerodynamic characteristics as a function of the shape is useful in exploring the full design space. Local inclination methods for aerodynamic analysis have proven sufficiently accurate for use at such a design stage, but manual selection of methods over the vehicle is inefficient for the exploration of a large number of design possibilities. This paper describes the model of an aerodynamic analysis code, written for use in conceptual vehicle shape optimization, which includes an automatic method selection algorithm. Panel shielding is also included in the analysis code to allow for the analysis of more complex geometries. The models used for the shape and aerodynamics are described and results for the Space Shuttle and Apollo are compared to wind tunnel data. They show an accuracy of better than 15% for most cases, which is sufficient for the use in conceptual design. Panel shielding is shown to be important in the prediction of control derivatives at low angle of attack, as well as the prediction of lateral stability derivatives. Finally, a simple guidance algorithm is used to assess the impact of the errors in the aerodynamic coefficients on the vehicle heat load and ground track length. Both show discrepancies of less than 10%.

Verified Interval Orbit Propagation in Satellite Collision Avoidance

Verified interval integration methods enclose a solution set corresponding to interval initial values and parameters, and bound integration and rounding errors. Verified methods suffer from overestimation of the solution, i.e., non-solutions are also included in the solution enclosure. Two verified integration methods, interval Taylor-series and Taylor-model based methods and their implementation in VNODE-LP and VSPODE, are used to reduce overestimation in verified satellite orbit propagation. Furthermore, two orbital state models based on integration constants, the modified equinoctial elements (MEE) and unified state model (USM), are used to reduce overestimation. Earth-satellite trajectories propagated using VSPODE have 2-3 times less overestimation than those propagated using VNODE-LP. Using the USM and MEE state models, overestimation is further reduced by a factor 4 to 10, depending on initial and parameter uncertainties. It is demonstrated that verified collision detection is feasible and may contribute to prevent satellite collisions to reduce future space debris.

Performance Aspects of Orbit Propagation using the Unified State Model

The Unified State Model is a method for expressing orbits using a set of seven elements. The elements consist of a quaternion and three parameters based on the velocity hodograph. The equations of this model and the background theory necessary to understand them have been shown here. Numerical simulations comparing the Unified State Model with the traditional Cartesian coordinates have been carried out for perturbed orbits, orbits with low-thrust propulsion, and a solar sailing trajectory. The Unified State Model outperforms Cartesian coordinates for all cases in terms of accuracy and computational speed, except for highly eccentric perturbed orbits. The performance of the Unified State Model is exceptionally better for the case of orbits with continuous low-thrust propulsion with CPU simulation time being an order of magnitude lower than for the simulation using Cartesian coordinates. This makes the Unified State Model especially suited for use in trajectory simulators and optimizers.

Shape optimisation for a small experimental re-entry module

For the development of reusable launchers, new technology has to be developed and tested that requires a hypersonic environment that cannot be reproduced in ground-based facilities. Small and low cost re-entry modules are needed for hypersonic experiments and the shape optimisation for such a re-entry module taking into account the required performance is presented in this paper. For the generic module shape, a blunted bi-cone has been selected that is simple to manufacture, has good stability properties and good potentials for various aerodynamic and material experiments. The cooling of the spherical nose is based on nucleate pool boiling of water. The aerodynamic design is done using a Response Surface Methodology (RSM) that includes the Taguchi method, a parametric multi-disciplinary design method with which the effects of changing several geometric design parameters in an ‘all-at-the-same-time’ approach can be studied, instead of the more traditional ‘one-at-atime’ approach. Each of the design iterations includes an aerodynamic analysis based on the Modified Newtonian method and a three-degrees-of-freedom trajectory analysis. Generating response surfaces for each of the performance indices and optimising them with a multiobjective optimisation method, a set of geometric parameters is found that gives the best alternative for each of the performance indices. Results of the verification of the (sub-) optimal design are included.