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This paper focuses on the attitude control and propellant slosh suppression of aeroelastic launch vehicles in a turbulent atmosphere. For a five-degree pitch-angle block command, the tracking performance of the selected Incremental Non-Linear Dynamic Inversion Sliding Mode Controller (INDI-SMC) shows excellent tracking performance. However, turbulence still inevitably leads to oscillatory behaviour in the swivel command. Various filter designs have been implemented to improve the smoothness of INDI-SMC. Using either a notch or band-pass filter in the sensor-feedback loops of pitch angle and pitch rate only marginally reduced the swivel oscillations, but did not solve the problem for the rigid-body control. For the flexible launcher with slosh dynamics, filtering of the sensor-feedback signals reduced the oscillations in swivel command, and elastic and slosh motion significantly, but could not completely remove them. The preliminary design of a rigid-body state observer has been included, and the results show that the INDI-SMC controller remains stable in the presence of engine dynamics, sloshing, flexible modes, input errors due to the use of rigid-body and slosh-state observers, while flying in a turbulent wind field.

Attitude control of conventional launchers is relatively easy and straightforward and gives an adequate performance when applied to the nominal vehicle and mission. However, in the presence of environmental disturbances and vehicle design uncertainties, more robust types of controllers are required to guarantee stable attitudes. This chapter discusses the application of Simple Adaptive Control for the pitch control of a conventional flexible launcher. Because of the large number of design parameters, an optimisation procedure based on an evolutionary algorithm has been applied. With a floating-point representation for the design parameters, stochastic universal sampling selection, arithmetic crossover and non-uniform mutation, the performance of the controller is analysed, and it is identified how the developed methodology can streamline the (conceptual) design phase. Application of Pareto ranking enabled the simultaneous minimisation of the state deviation and the control effort, while the oscillation of the control has been used as an optimisation criterion. A conclusive simulation shows the controller performance for the flexible launch system.

1. Introduction With the increasing interest in the Solar System's smaller bodies, quite a few missions have been sent to comets and asteroids, and more will be send in the near future. Due to the large distances involved, communication to command mission parameters takes a long time, which has a negative impact on operational safety. Autonomous navigation would be one of the key technologies that can make asteroid missions more robust, safe, and cost effective. This is especially true if one considers the unknown flight environment when the spacecraft is first encountering the body. Most asteroids and comets have a very irregular shape and unknown mass distribution. Therefore, knowledge about its irregular gravity field will be directly beneficial as input to orbital corrections and manoeuvre planning. 2. Methodology Based on a detailed ``real-world'' simulator, orbits around a small body, here the asteroid Eros-433, can be simulated that are perturbed by non-linear gravity effects, as well as solar-radiation pressure and third-body perturbations. Measuring land marks on the small-body's surface, processed (in a preliminary mission phase with Earth in the loop) to a detailed land-mark database, can be subsequently used to determine the (relative) position and velocity of the spacecraft with respect to the body. A next step would be to use measurements of orbital perturbations to estimate as many components of the body's gravity field as possible. Detailed knowledge about the gravity field can then be used for planning science and mission operations, such as orbital corrections. An in-house code is used to determine the spherical-harmonics coefficients Cnm and Snm, assuming an average asteroid density of 2.6212 g/cm3. Part of the navigation system are simplified sensors, which provide noisy position and/or velocity measurements. Interfacing with more realistic sensors is possible, but left as future work. The core of the navigation system is an augmented state estimator, based on the unscented Kalman filter for its applicability to highly non-linear systems. 3. Results Eros-433 is an elongated body, roughly resembling an ellipsoid, with dimensions of 33 by 13 by 13 km, and rotates with an approximate constant rate in 5.27 hours. For the ``real-world'' asteroid model, a gravity field up to order and degree 16 has been determined, see Figure 1. The results shown are the deviations from the point-mass model and thus highlighting the irregularities of Eros' gravity field. With increasing altitude these irregularities are much more ``smeared out'' and will thus be harder to estimate. The first simulations that were done aimed at estimating the central-field parameter, μ, and the first zonal harmonic, J2 = C2,0. To do so, several orbital altitudes ranging from 1500 down to 150 km were chosen. In terms of measurements data, it was assumed that only noisy position and velocity data were available at a frequency of 1 Hz. Figure 2 shows the estimates of μ. The remaining error is well below 0.1% for the different altitudes. With the estimated value of μ substituted in the filter, the estimate of J2 is seen to converge best at an altitude of 45 km: at higher altitudes the J2 effect starts to decrease and becomes less noticeable, whereas at lower altitudes the effect of the higher-order terms start dominating. The remaining error is about 6%. The last part of the research focusses on the estimation of higher-order coefficients. Idealised noisy position (noise level 10 m, which would be possible with current sensor technology) and velocity measurements are used, supported by ideal range measurements at 2 Hz. Figure 3a displays the errors of the full 8 degree and order coefficients for the estimation at an orbital height of 50 km. The coefficients of degree 6 and higher have significant errors, but the coefficients up to degree and order 5 can be estimated within errors of less than 30%. However, once μ and J2 have been estimated and subsequently put to constant values in the estimator, the estimation accuracy significantly improves, but only when the orbit is lowered to 35 km and the measurement time is extended to 10 days (Figure 3b). The average error is about 10%, with a few outliers. These errors are quite small, because of the relatively ideal measurements. Follow-on research should therefore focus on realistic sensor modelling and including instrument errors to the augmented state.

Verified interval orbit propagation provides mathematically guaranteed solutions of satellite position and velocity over time. These verified solutions are useful for conjunction analysis and other space-situational-awareness activities. Unfortunately, verified methods suffer from overestimation and explosive interval growth, limiting the possible propagation time and thus their applicability. Different orbital-element state models have been shown to increase the maximum propagation time to a degree, but at the expense of significant overestimation introduced by the state transformations. This paper proposes the Dromo state model for verified integration. Dromo is a regularized variation-of-parameter formulation of the perturbed two-body equations of motion. Taylor models are implemented for both integration and transformation. Moreover, a technique is developed for dealing with time uncertainty resulting from verified regularized propagation. Dromo significantly prolongs the maximum forecasting window, producing verified trajectories of days up to weeks in duration for the low Earth orbit regime. A sensitivity analysis of the integrator settings identifies combinations that produce stable and computationally efficient solutions. A sensitivity study of the orbital parameters shows that the method is applicable to a large orbital regime, especially for low Earth orbit regions that contain high densities of space debris.

The beginning of the conceptual design phase of (re)entry missions requires aerodynamic methods to reduce the initial design space. For this purpose, full computational fluid dynamics (CFD) simulations are unsuitable due to their computational requirements. Rapid hypersonic methods are thus often employed to approximate the heat flux and skin friction on the most critical parts of the (re)entry vehicle, such as the nose and the leading edges. However, the vast majority of these rapid methods only allow for a computation of these parameters at specific fixed locations and not on the other parts of the vehicle. One method that overcomes this is the axisymmetric analogue method, that determines the entire viscous flowfield from the inviscid flowfield solution. This method has typically been coupled to inviscid Euler simulations, but even Euler simulations can still consume a lot of computational time. In earlier research, it was shown that a reasonable accuracy can also be obtained if this method is coupled with the inviscid flowfield computed via the modified Newtonian technique. In this paper, we extend the validation and estimation of the uncertainties of this method using CFD, evaluate the respective corrections for thermal and chemical fluxes separately, and apply these corrections back to the solver. The biconic DART vehicle, partial optimisation of which was presented in the previous paper, is revisited, here optimising only four parameters instead of five as originally intended, as using five parameters resulted in an unfeasible geometry. We perform a full response surface methodology and analysis of variance accounting for the CFD corrections and examine the final optimised design also again with the Newtonian/axisymmetric code. The proposed methodology leads to a small underestimate of the heat fluxes, but is considered sufficient for the conceptual design phase.

We present a 3 Degrees of Freedom mission design and analysis for in-situ probing of Uranus' atmosphere consisting of two un-propelled gliders and one orbiting spacecraft in continuous line of sight. We focus on the study of the gliders' navigation and science modules. Because of the lack of a Global Navigation Satellite System around Uranus and the ineffective use of optical sensors due to the planet's large distance to the Sun and high atmospheric opacity, the post-processing relation between the vehicles' estimated state and measured scientific data is investigated to yield accurate state estimations. In-situ probing by the two gliders will make it possible to measure spatially variable atmospheric properties over a flight duration of up to 12 Earth days, as compared to a few hours for a conventional descent probe. Future work will include a 6 Degrees of Freedom simulation of the vehicles' flight, the chosen planet's wind model, a Flush Air Data Sensor as an additional navigation sensor, and a band-pass filter to reduce the estimated variables' noise.

Re-entry shape and trajectory optimisation studies typically require hundreds to thousands of flow solutions to resolve the heat transfer and skin friction. Due to the fact that full CFD simulations and even Euler simulations are typically very expensive, this work presents a development of a much simpler technique, in which the modified Newtonian approximation is coupled with the axisymmetric-analogue-based viscous method. Afterwards, the developed technique is applied to the shape optimisation study of the DART module originally conducted by Sudmeijer and Mooij in 2002. Large differences in heat fluxes are observed mainly owing to the fact that the present method estimates the transition region from approximate transition criteria while the original study assumes the transition point to be fixed at the interface between the cone and the flare. This emphasises the sensitivity of design optimisation studies on delicate parameters such as the location of transition, which typically cannot be accurately predicted during the conceptual phase. Finally, the limitations of the current method are identified and the areas in which the technique will be improved further are outlined.

Precision landing is an anticipated technology for future interplanetary missions. Autonomous spacecraft Entry, Descent and Landing (EDL) on the surface of a planetary body with a degree of precision in the order of meters is highly challenging. In this paper, a successive convexification guidance algorithm is utilized to simulate autonomous precision landing sequences on Saturn’s moon Titan. Due to its unique geophysical features, studying the science of matter within Titan’s atmosphere and beneath its surface is one of NASA’s most important planetary science objectives. As part of the Space Exploration Technology Directorate, a parafoil is proposed for landing on Titan due to its cost effectiveness, ease of deployment, low mass compared to the prospective payload and capabilities of precise autonomous delivery. This paper focuses on path optimization and guidance law development for high-fidelity dynamics parafoil tuning in the dense and adverse wind atmosphere of Titan, defined as a nonlinear and nonconvex optimal control problem. The powerful successive convexification method is used to solve the problem accordingly. The algorithm is designed such that the converged solution adheres to the nonlinear dynamics and kinematics in accordance with the original formulation, while respecting the state and control constraints. The six-degree-of-freedom (6DoF) simulations results show that this robust method is suitable for autonomous interplanetary applications.

This paper proposes a two-stage optimization approach for aerocapture guidance. In classical entry guidance systems, deterministic optimization is used. Large-scale and short-scale density perturbations may strongly affect the performance of the guidance system, and variations in those are usually not accounted for when computing the command. In this work, perturbations that affect the trajectory at future time-steps are taken into consideration when computing the commanded bank angle. The chosen command is optimal based on a set of possible future perturbations, after the observation of which, a correction can be made. Both two-stage stochastic and two-stage robust optimization are proposed as a solution. In a Monte Carlo analysis consisting of 50 runs, the two-stage robust optimization guidance outperforms an optimal, deterministic, numeric predictor-corrector guidance. Excluding one outlier, also the two-stage stochastic optimization makes the guidance perform better than an optimal deterministic numeric predictor-corrector. With either approach, the computational demands are increased by about thirty times compared to an optimal numeric predictor-corrector. Much of the computation time increase may be reduced by parallelization. On the other hand, the extensive tuning required for the optimal numeric predictor-corrector is not needed for the two-stage optimization guidance, making this approach conceptually more robust. Better modeling of the environment may help further improve the performance. Finally, an approximation to the two-stage robust optimization approach is developed. The guidance has computational requirements only four times larger than those of the optimal numeric predictor-corrector guidance, but can be parallelized into two threads, and, except for two outliers, it offers improved performance.

This paper focuses on the attitude control and propellant slosh suppression of aeroelastic launch vehicles. Four candidate controllers are proposed: the Linear Quadratic Regulator (LQR), the Incremental Non-linear Dynamic Inversion (INDI) control, the Incremental Sliding Mode Control (INDI-SMC), and the Feedback Linearisation-based Sliding Mode Control (FL-SMC). Theoretical analyses show INDI itself is unable to deal with under-actuated systems. Therefore, when applied to the launch vehicle directly, it cannot simultaneously track the pitch command and effectively suppress the slosh dynamics. This issue is solved by INDI-SMC, which also has enhanced robustness against both matched and unmatched uncertainties. Furthermore, despite its reduced model dependency, INDI-SMC has better robustness against model uncertainties and external disturbances than FL-SMC. These merits of INDI-SMC are verified by various simulation results. First, when the nominal plant configuration is adopted, the system using INDI-SMC has the smallest pitch-angle tracking error. The slosh motion is also effectively damped out. Second, Monte-Carlo studies are used to test the robustness of LQR, INDI-SMC, and FL-SMC to parametric uncertainties. Among these three controllers, LQR shows the worst performance and largest control-effort outliers. On the contrary, both INDI-SMC and FL-SMC can resist a wider range of perturbations without significant performance degradation. Even so, the tracking and slosh damping performance of INDI-SMC is still the best. Finally, both INDI-SMC and FL-SMC show robustness against unmodeled dynamics, while the robust performance of INDI-SMC is superior.

Long and slender bodies, such as (small) conventional launch systems, may suffer from an unwanted coupling between the rigid body and its flexible modes. The current research treats the launch vehicle as a flexible beam with lumped masses to account for the subsystems and fuel, using a three-dimensional assumed-modes method with longitudinal and lateral effects. Given the response of a simple proportional-derivative controller as benchmark, the performance of an Incremental Non-linear Dynamic Inversion (INDI) controller and a system based on Simple Adaptive Control is studied for a number of distinct manoeuvres. Of the three, the INDI controller has a superior performance and is not affected much by the inclusion of engine dynamics and flexible modes. Properly designed filters in the feedback loops show that the rigid-body response can be decoupled from the flexible-body motion, although controller gains need to be adapted to this new configuration. A second advantage is that structural vibration is reduced. Finally, the inclusion of gyroscopes, placed far away from the launcher's centre of mass and which suffer from the effects of aeroelasticity, does not lead to a large performance degradation if both the pitch-rate and pitch-angle feedback signals are properly filtered. To just counter the effects of flexibility, band-pass filters are best suited. However, the effect of flexibility on the gyroscope output is best compensated by notch filters.

This paper examines the dynamics and control involved in the active removal of a large space debris - Envisat. In the light of European Space Agaency's e.deorbit mission, the current mission scenario consists of a chaser satellite, synchronising, capturing and detumbling the large uncooperative debris. A unique multibody approach based on floating frames is used to model the Flexible Multibody Dynamics. The controllability characteristics of a linear PD controller and an Incremental Nonlinear Dynamic Inversion controller are also studied. From previous research, it was found that in the docked configuration, the control of the system could not be achieved due to the complex elastic dynamics originating from the large solar panel of Envisat (14.2~m). This paper focuses on achieving robust control of this phase through advanced design choices in both structural and control aspects. It was found that the inclusion of structural damping and reduction of control limits, lead to a major improvement in controllability of the system. A basic digital notch filter did not prove effective for more robust control. This is because only a high order filter could account for the low frequency vibrations. Further, it also required retuning of the controller within the flexible model. These two aspects proved to be beyond computational capabilities. A simple control strategy of introducing a dead-band in control loop improved the control response considerably by allowing the disturbing frequencies to dampen out.

This paper focusses on the design of a sliding-mode observer to estimate the states of the slosh-mass position and velocity. These states are subsequently used as feedback signals for the attitude controller. A simple proportional-derivative (PD) rigid-body controller cannot safely execute attitude commands without encountering stability problems, and even in a steady state situation a non-zero slosh-mass positions leads to jittering in the swivel angle. With slosh-state feedback, even a non-optimal PD controller exhibits a smooth response with limited actuator activity. However, a slight uncertainty in the slosh-model parameters of the observer will negatively affect the controller performance, and, as a consequence, increase the amplitude and oscillatory behaviour of the liquids. An additional observer will be required to estimate the correct parameter values. A more robust incremental non-linear dynamic inversion controller improves the controller response, but further development will be required, for instance, by including notch filters in the feedback loops.

By simulating full lift-up and full lift-down trajectories, the limit initial conditions for which successful aerocapture is possible have been determined. These boundaries form the entry corridor for aerocapture, and have been developed for Earth, Mars, and Venus. Moreover, two different vehicles representative for either manned missions or sample return were studied. The vehicle configurations were varied to study the effect of mass and lift on the size of the entry corridor. From this investigation it was determined that increasing the lift produced by the vehicle is a more effective method of widening the entry corridor than decreasing the mass. Furthermore, it was found that a lift-up-lift-down bang-bang type of trajectory minimises both the total $\Delta V$, the load factor, and the heat load. An optimisation scheme was set up to determine the optimal switch time, which showed that for nearly all initial conditions that fall within the boundaries specified by the developed entry corridors, optimal aerocapture could indeed be achieved with an apoapsis correcting manoeuvre with a $\Delta V$ smaller than 0.01 m/s.

Satellite reentry predictions are used to determine the time and location of impacts of decaying objects. These predictions are complicated by uncertainties in the initial state and environment mod...

This paper investigates the efficacy of dual quaternion filtering in the realm of asteroid missions. The main contribution is the development of a dual quaternion relative navigation filter applied...

Accurate knowledge of satellite orbit errors is essential for many types of analyses. Unfortunately, for two-line elements (TLEs) this is not available. This paper presents a weighted differencing method using robust least-squares regression for estimating many important error characteristics. The method is applied to both classic and enhanced TLEs, compared to previous implementations, and validated using Global Positioning System (GPS) solutions for the GOCE satellite in Low-Earth Orbit (LEO), prior to its re-entry. The method is found to be more accurate than previous TLE differencing efforts in estimating initial uncertainty, as well as error growth. The method also proves more reliable and requires no data filtering (such as outlier removal). Sensitivity analysis shows a strong relationship between argument of latitude and covariance (standard deviations and correlations), which the method is able to approximate. Overall, the method proves accurate, computationally fast, and robust, and is applicable to any object in the satellite catalogue (SATCAT).

In this paper, an improved predictor-corrector algorithm with hybrid lateral logic is proposed to enhance the maneuverability during re-entry, which lets the reusable launch vehicle (RLV) possess the ability of avoiding a no-fly zone. First, the longitudinal guidance is converted to the quasi-equilibrium guidance algorithm once current states satisfy the joint point condition. During each guidance period in the gliding phase, a quadratic parametric model is devised to adjust the magnitude of the bank angle instead of utilizing the linear one. Then, an artificial potential field based lateral guidance law is designed for no-fly zone avoidance by transforming the problem into finding the reference heading angle. But the bank angle reverses while the RLV is approaching the target, which may aggravate the instability problem and increase the energy consumption. Aiming at above problem, the lateral guidance is combined with conventional heading angle deadband corridor, which works when the RLV leaves the influence area of the no-fly zone. Finally, numerical simulations show that the proposed lateral guidance logic not only is effective for no-fly zone, but also performs well in reducing the times of reversals. The Monte Carlo simulation results further demonstrate the robustness of the above guidance algorithm considering the random initial dispersions and errors.

This paper proposes a real-time implementable Lyapunov-based guidance algorithm for small-body proximity operations. The guidance solution accounts for the secular effects coming from the oblateness of the main body and for the non-impact condition between the latter one and the spacecraft. The first perturbation with respect to the two-body problem is implemented via an update of the Keplerian elements based on the secular precessing ellipse model, while the second one through a check on the radial distance at periapsis. Advantages are taken of the fact that real-time guidance algorithms require only the knowledge of the current state of the spacecraft and of the target. The proposed mission scenario implements an approximate model of 433 Eros as the main body and a low-thrust point mass spacecraft executing orbital transfers about it. It is shown that, for transfers lasting tens of days, ignoring the oblateness of the main body results in errors on the order of ten degrees for the right ascension of the ascending node, the mean anomaly, and the argument of periapsis. These errors are reduced by two orders of magnitude when the proposed guidance algorithm is applied.

A key requirement for enabling manned Mars missions is the capability to land a payload within 100 m from its nominal target. State of the art technology, however, is still far from reaching this objective. A possible solution for improving the landing accuracy on Mars is the development of a guidance system that exploits thrusters installed on the backshell of the descent vehicle to control its position during the parachute descent. The research summarised here aims at evaluating how the introduction of the guidance system influences the stability properties of the parachute-payload spacecraft and what performance the resulting closed-loop system can achieve. To do this, first the dynamics of the vehicle, modelled as a rigid body, has been investigated analytically and then its response to different external perturbations was simulated by means of a multibody model. The results are encouraging and prove that the parachute descent guidance system could efficiently contribute to improving the landing accuracy on Mars.

The development of fully reusable launch systems has been the topic of many studies since the 1960s. Over the years, several aspects of both so-called single- and two-stage-to-orbit space planes have provided many interesting research topics. Amongst others, the constrained trajectory optimisation has proven to be a challenging subject. In this chapter, an inverse-dynamics approach is combined with trajectory optimisation and analysis, by discretising a representative (vertical-plane) ascent trajectory into a number of flight segments, and by parametrising the guidance in terms of flight-path angle as a function of altitude. When the individual guidance parameters are varied, the effect on performance indices payload mass and integrated heat load can be analysed. This can subsequently lead to a refinement of the trajectory. To do so with limited effort, design-of-experiment techniques are used. It is shown that with this relatively simple simulation scheme, combined with variance analysis and response-surface methodology, the insight in the trajectory dynamics can be increased. Alternatively, this method can be used as refinement to an otherwise (local) optimum trajectory. It is stressed, though, that the application of design of experiments to the ascent-trajectory problem cannot replace numerical optimisation. Finally, the impact of using thrust-vector control as a means to (partially) trim the vehicle shows significant fuel savings and should therefore be included in the optimisation process.

Examples of control problems occurring during flight tests of fighter aircraft are well documented. In many cases, the cause of the problem could be characterised as inadequate modelling or other inappropriate treatment of the aeroelastic effects on the vehicle dynamics and/or the flight-control design. However, such problems are not restricted to just aircraft. Especially long and slender bodies, such as (small) conventional launch systems, may suffer from an unwanted coupling between the rigid body and its flexible modes. In addition, due to fuel consumption during the flight, the change in atmospheric environ-ment, and aerodynamic effects, the entire flight profile should be examined, rather than a single worst-case point, to identify the stability and controllability performance of the launch vehicle. The current research treats the launch vehicle as a flexible beam with lumped masses to account for the subsystems and fuel, using a three-dimensional assumed-modes method with longitudinal and lateral effects. Sloshing is added to the system as a spring-mass-damper, where the liquid oxidiser and fuel are modelled separately. Multiple discrete flight points are considered for aeroelastic analysis at different stages of the ascent profile. These individual points are analysed as quasi-static and quasi-steady inputs for a continuous flight trajectory simulation. The stability and controllability for the entire flight under the influence of wind gust and turbulence is estimated based on this simulation. While simulating the manoeuvres with the sloshing model included it became apparent that the (very) simple PD controller was not able to stabilise the vehicle. Results concerning the influence of sloshing can therefore not be shown. The development of a more robust adaptive controller is currently on its way, and it is expected that this non-linear controller will be better suited to handle the strong non-linearities of the combination of flexibility, sloshing and a demanding wind environment.

Spaceplanes represent a new promising concept for space flight. A spaceplane is a reusable, safe, efficient and economical space transportation system that can generate lift during its atmospheric flight, analogously to an aircraft, and it is able to travel in space as a spacecraft. What makes spaceplanes so attractive is the possibility of reusing the system for more than one mission, and the flexibility that they allow in mission. The growing interest in hypersonic spaceplanes requires that these vehicles have adequate properties of safety of flight and ease of controllability in nominal and off-nominal conditions. From this it follows the need to study their flying qualities and controllability characteristics. The current thesis addresses the flying-quality and controllability analyses, together with the development of a control system capable of improving these properties. The analyses are conducted along the ascent and re-entry trajectory of a representative single-stage-to-orbit spaceplane, the Festip System Study concept 1. The stability, trim capabilities and flying qualities of the open-loop system are analysed by studying the eigenvalues and eigenvectors of the state matrix of the state-space model. Comparing the obtained results with the requirements specified in the military documents MIL-F-8785C and MIL-HDBK-1797 for subsonic vehicles, it is possible to conclude that the reference vehicle is dynamically unstable. Thus, the need for an advanced control system arises. One concept seems particularly interesting for this application: the adaptive control system, which is characterised by a low sensitivity to disturbances thanks to its adaptive gains. Not only the control system is design to be optimal in terms of integrated control error and effort, but also a robust design methodology is applied to identify a control design that is as insensitive as possible to uncertainties of the input and design parameters. The responses for both longitudinal and lateral control in nominal and off-nominal conditions are simulated and evaluated. It results that the system behaviour is strongly related to the control system performance. The robust and advanced control system is able to stabilise the vehicle with relatively low control effort and minimise the effect of disturbances, guaranteeing safety of flight and mission success.

Next-generation planetary tracking methods, such as interplanetary laser ranging (ILR) and same-beam interferometry (SBI) promise an orders-of-magnitude increase in the accuracy of measurements of solar system dynamics. This requires a reconsideration of modelling strategies for the translational and rotational dynamics of natural bodies, to ensure that model errors are well below the measurement uncertainties. The influence of the gravitational interaction of the full mass distributions of celestial bodies, the so-called figure-figure effects, will need to be included for selected future missions. The mathematical formulation of this problem to arbitrary degree is often provided in an elegant and compact manner that is not trivially relatable to the formulation used in space geodesy and ephemeris generation. This complicates the robust implementation of such a model in operational software packages. We formulate the problem in a manner that is directly compatible with the implementation used in typical dynamical modelling codes: in terms of spherical harmonic coefficients and Legendre polynomials. An analytical formulation for the associated variational equations for both translational and rotational motion is derived. We apply our methodology to both Phobos and the KW4 binary asteroid system, to analyze the influence of figure-figure effects during estimation from next-generation tracking data. For the case of Phobos, omitting these effects during estimation results in relative errors of $0.42\%$ and $0.065\%$ for the $\bar{C}_{20}$ and $\bar{C}_{22}$ spherical harmonic gravity field coefficients, respectively. These values are below current uncertainties, but orders of magnitude larger than those obtained from past simulations for accurate tracking of a future Phobos lander., 29 pages, 4 figures (accepted for publication by Astrophysics and Space Science)

The space environment is ever-changing with space structures getting larger and the orbits getting increasingly crowded with time. This creates a need for removal of large defunct satellites to avoid the disastrous Kessler syndrome, which poses a major threat to the future of space exploration. This research examines the dynamics and control involved in the active removal of a large space debris - Envisat. European Space Agency's e.deorbit mission aims to de-orbit Envisat using a chaser satellite, which synchronises, docks, detumbles and deorbits it. The presence of large flexible appendages makes the configuration prone to elastic perturbations leading to complex dynamics that cannot be represented using rigid body dynamics. Therefore, a unique multibody approach based on the absolute interface coordinates in the floating frame formulation is used to model the Flexible Multibody Dynamics. The novel method proves to provide a good balance between computation time and efficiency for the control application. The controllability characteristics of two phases of the e.deorbit mission are analysed using a linear PD controller and an Incremental Nonlinear Dynamic Inversion controller. For the first phase, both controllers successfully synchronise the chaser with the target debris tumbling at the rate of 3.5 deg/s about all axes. However, during the detumbling phase, the large appendage (14.2 m) in the stacked configuration introduces complex dynamics, which could not be stabilised completely by applied controllers. Nonetheless, interesting relationships could be established between the dynamics and control of the system, which will facilitate robust control design in future work.

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 ...

For future landings on any celestial body, increasing the lander autonomy as well as decreasing risk are primary objectives. Both risk reduction and an increase in autonomy can be achieved by including hazard detection and avoidance in the guidance, navigation, and control loop. One of the main challenges in hazard detection and avoidance is the reconstruction of accurate elevation models, as well as slope and roughness maps. Multiple methods for acquiring the inputs for hazard maps are available. The main distinction can be made between active and passive methods. Passive methods (cameras) have budgetary advantages compared to active sensors (radar, light detection and ranging). However, it is necessary to proof that these methods deliver sufficiently good maps. Therefore, this paper discusses hazard detection using stereo vision. To facilitate a successful landing not more than 1% wrong detections (hazards that are not identified) are allowed. Based on a sensitivity analysis it was found that using a stereo set-up at a baseline of ≤ 2 m is feasible at altitudes of ≤ 200 m defining false positives of less than 1%. It was thus shown that stereo-based hazard detection is an effective means to decrease the landing risk and increase the lander autonomy. In conclusion, the proposed algorithm is a promising candidate for future landers.

The aim of this research is to investigate a debris-remediation technique where a chaser performs a rendezvous with the debris, establishes a rigid-link connection, and actively de-orbits the debris. ESA׳s satellite Envisat has been used as a design case. The research assessed passive safety aspects of the final-approach manoeuvres by analysing the resulting trajectories after thrust inhibit. Next, the research explored the possibility for continuous ground communication by considering the chain of European space tracking (ESTRACK) ground stations (located mainly in Europe). Furthermore, obstruction of the communication signal by the target was studied. Last, the research studies the illumination conditions encountered by the chaser, where obscuration of the Sun by the target was taken into account. Each of these elements are studied for the final approach only. In the topic of passive safety, the results confirm that manoeuvres on H-bar are passively unsafe, and indicate this also for the fly-around manoeuvres along the natural orbital motion. It can be concluded from the communication analysis that the maximum duration of the uninterrupted window varies between 22 and 32 min, using the chain of core ESTRACK ground stations. However, the study on communication blockage shows that frequent communication gaps can occur, with the longest gaps being in the order of one minute in duration. In the field of illumination, it can be concluded that correct target illumination and sensor visibility cannot be guaranteed. Furthermore, the average solar-array area available during final approach varies between 35% and 75%, due to both incorrect pointing of the solar array by the chaser and obscuration by the target.

Aiming at unknown disturbances/uncertainties, partial effectiveness loss fault (PELF) and stuck failure (SF) of the actuator, a composite robust fault-tolerant control strategy based on incremental backstepping (IBS) is proposed for a reusable launch vehicle (RLV) during re-entry. By converting PELF to disturbances/uncertainties, this paper presents an incremental form of disturbance observer based on an improved inverse hyperbolic sine tracking differentiator (IHSTD) to compensate these interference terms originally ignored in the IBS design process. Furthermore, a failure symbol matrix is set to control the on-off states of the reaction control system of the RLV to make up for the missing torque of the actuator SF, which can strengthen the fault-tolerance capability of the control system. The simulation results show that the tracking effect of the proposed method on the attitude-angle commands is better than traditional backstepping with disturbance observer, and the presented control allocation strategy is capable of timely resolving the actuator SF problem to ensure stability of flight.

This paper gives an overview about the Aurora reusable space launcher concept study that was initiated in late-2015/early-2016. Within the Aurora study, several spaceplane-like vehicle configurations with different geometries, propulsion systems and mission profiles will be designed, investigated and evaluated with respect to their technical and economic feasibility. The first part of this paper will discuss the study logic and the current status of the Aurora studies and introduces the first vehicle configurations and their system design status. As the identification of highly efficient structural designs is of particular interest for Aurora, the structural design and analysis approach will be discussed in higher level of detail. A special design feature of the Aurora vehicle configurations is the utilization of the novel thin-ply composite material technology for structural mass reductions. Therefore, the second part of this paper will briefly discuss this technology and investigate the application and potential mass savings on vehicle level within simplified structural analysis studies. The results indicate that significant mass savings could be possible. Finally, an outlook on the next steps is provided.

To maximize the payload mass, an aerocapture trajectory should be flown in such a way that both the final ΔV and the total heat load are minimized. At very high velocities, the heating due to radiation of high temperature gases in the shock-layer exceeds the heat due to convection. For some aerocapture missions, such a heat load can be 15 times larger than the heat load due to convection. Thus, convective heat load may in some cases be neglected. It is analytically proven that radiative heat load is minimized by the same trajectory that Minimizes The Final Δ V: a bang-bang trajectory, with full lift-up, full lift-down commands. Next, a novel guidance that plans a bang-bang trajectory with constraints in the attitude kinematics is introduced. This allows for achieving an optimal trajectory with only one parameter to be tuned. For the case studied, values of ΔV as low as 100 m/s can be ensured for entry angles between -6° and -5° and a large spectrum of perturbations; with the same guidance, radiative heat load is reduced by up to 20% with respect to traditional aerocapture-guidance methods. Finally, a lateral guidance that makes use of information on the final inclination of the predicted trajectory is introduced. Such guidance allows for very high accuracy in the inclination requirements with only two reversals, and also requires only a single parameter to be tuned. Depending on the tuning, a maximum inclination error of less than 0.1° can be guaranteed.

This paper presents a novel reentry guidance architecture that aims to improve cur- rent mission safety by enabling zero-wait orbital aborts. To do so, an on-board trajectory planner based on Adaptive Multivariate Pseudospectral Interpolation (AMPI) is devel- oped. This planner generates mathematical approximations of the reference trajectory to be tracked by comparing the vehicle-state information at the entry point against a pre-computed trajectory database that is stored on-board. Such a trajectory database is populated by systematically solving a number of optimal-control problems, which ulti- mately generate a set of reference angle-of-attack and bank-angle profiles that feed a Linear Quadratic Regulator trajectory tracker. The architecture proposed in this paper is tested using a scaled-down version of the X-38, SPHYNX, a small lifting-body reentry vehicle developed by the European Space Agency. The proposed on-board trajectory planner is able to generate reference commands in the order of a few milliseconds, thus proving the ability to run the system in current ight-hardware. Furthermore, a Monte Carlo test campaign showed that the trajectory database can be compressed with a loss-less process without having a detrimental impact on the landing dispersions and constraint compliance, achieving database sizes as low as 18 MB. Finally, the presented architecture allows for easy targeting of any landing site, as long as reference guidance commands are stored in the database.

Simple adaptive control can be applied to any non-linear system, provided it is almost strictly passive. A necessary condition in that case is that each control channel is single-input single-output. Re-entry vehicles, where reaction and aerodynamic control are combined, do not fulfill that condition. In particular, gliding vehicles at large angle of attack, where the vertical stabiliser and/or winglets are partially shielded by the fuselage, need yaw thrusters to support the weak rudder control. Instead of commanding the actuators directly, a required moment per axis should therefore be commanded, and the actual actuator allocation should be done outside the controller. Not only does this make the closed-loop system almost strictly passive, but also the controller design is simplified. The current paper addresses such a control-system design, and shows response results for both longitudinal and lateral control in several mission phases that require the use of thrusters, control surfaces, and a combination of both. With relatively low design effort, a smooth and quick response is obtained that can easily be extended to integrated attitude control along the complete re-entry trajectory.

In this paper a new drag-energy scheme for atmospheric entry guidance, based on the use of pseudospectral methods and convex optimization, is proposed. One of the most successful technologies to deal with atmospheric entry is the class of drag-tracking schemes, a direct heritage of the Space Shuttle program. The method that we propose exploits the drag-dynamics, and allows for an efficient automatic design of an optimal entry profile satisfying all the longitudinal constraints acting on the vehicle. A new representation of the entry-guidance problem, able to loss-less convexify the formulation, is provided. Numerical simulations confirm the validity of the proposed scheme as tool for further improving the autonomy of modern entry guidance systems, with a mean final range-to-go error at the end of entry phase smaller than 2 km, and the capability to re-compute a complete constrained trajectory to meet the mission requirements.

To limit the mass of the vehicle’s thermal protection system, an optimal trajectory that minimises the total integrated heat load should be own. This means that the maxi- mum heat- ux constraint is followed for as long as possible, until the maximum mechanical load is encountered. Flying as close to this load as possible contributes to minimising the heat load as well. The guidance system to track the path constraints includes three components: a semi-analytical guidance that produces nominal bank-angle commands, an inner-loop tracking system based on linearised output feedback, and an outer-loop system based on simple adaptive control to remove the remaining error. The former two guidance components have been taken from an earlier design, and the performance gain after adding the adaptive system is studied in detail. The ight system under consideration is a hyper- sonic test vehicle of which the stagnation heat- ux should not exceed 1,700 kW/m2, with a limit of the mechanical load of 5g. The results show that the adaptive tracking system extends the duration of heat- ux tracking and is able to track tighter at even less guidance effort, but reduces the fiight range because of that. A simultaneous optimisation of these two con icting objectives should be pursued to reFIne the guidance-system design in case both have requirements to be met. In none of the cases considered, the g-load constraint was violated, although a more detailed analysis is required to make this part of the guid- ance more robust. In all other aspects, the combination of an inner-loop and outer-loop tracking has made the path-constraint tracking more robust and keeps the errors below acceptable levels.

Hazard-detection and avoidance systems will become an important asset for next-generation landing and exploration missions. To date, multiple studies into these systems were conducted to develop the methods needed and to demonstrate their performance in mainly software-based tests. Few studies were able to demonstrate the performance of the algorithm in hardware-in-the-loop tests, as these are usually difficult to set up and expensive to execute. In this paper, the hardware-in-the-loop testing of a stereo-vision based hazard detection algorithm is presented. It was performed with the Testbed for Robotic Optical Navigation (TRON) at the German Aerospace Center (DLR) in Bremen, Germany. Since this testbed only allows for testing in a scaled environment, one of the challenging tasks was to design a scaled-down test set-up to represent a real-life lunar descent. The hardware-in-the-loop testing confirmed the results obtained during the earlier software-in-the-loop testing, that stereo vision can successfully be used for hazard detection during planetary descent.

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...

One of the most powerful analysis tools to deal with entry-guidance problems is the possibility to formulate them as optimal control problems. Environmental constraints, actuator limits, and strict requirements on the final conditions can be efficiently transcribed, resulting in a discrete, finite-dimension nonlinear programming (NLP) problem. However, NLP problems require a computational power, which often exceeds the vehicle’s onboard capabilities. Moreover, it is important to ensure that the solution can be adapted to the actual flight conditions, which can differ from the nominal scenario. This paper proposes an approach based on an efficient use of multivariate pseudospectral interpolation scheme to generate real-time capable entry guidance solutions. The proposed onboard trajectory generation algorithm is able to deal with wide dispersions at the entry interface, and can improve the lateral performance in cases where the classic bank-reversal logic is not sufficient. The interpolation is applied to subspaces of a database of pre-computed trajectories, which can be efficiently stored onboard. The method is here proposed for initial-conditions variations, but can be applied to every mission parameter, which allows to find a corresponding optimal solution. Results have been generated for SHEFEX-3, an entry demonstrator vehicle, which was planned by the German Aerospace Center. Monte-Carlo simulations show how this approach is applicable, and yields significant improvements both in longitudinal and lateral guidance performance, with an improvement of the dispersion area of about 96%.

In the design of a Model Reference Adaptive Control system, a reference model serves as the (well-known) basis through which system and user requirements can find their way into the design. By tuning the design parameters, the response of the actual vehicle should track the response of the reference model as well as possible. In the current study, the actual vehicle is an unpowered, winged re-entry vehicle similar to the Space Shuttle. To understand the influence of each of the design parameters on the controller performance, a sensitivity analysis is conducted on sample frequency, weighting-matrix coefficients and compensator parameters. Also the influence of varying flight conditions is studied. Normalisation of the reference signals has led to a more unified definition of the design parameters, and has introduced more independence from the flight conditions. Even though the influence of the design parameters is well understood, the large number makes tuning them a tedious process. An automated optimisation procedure should be used, since the interaction between the related design parameters makes an all-in-one-go approach mandatory. Experience of the designer, though, can easily reduce the design time, if a sub-optimal yet satisfactory response is strived for.