Limited Thrust Research Articles

A novel concept of cost-effective active debris removal spacecraft system

Post Mission Disposal (PMD) alone is an insufficient solution to stabilize the orbital debris environment, but combining Active Debris Removal (ADR) with PMD and other means of debris mitigation is considered a promising remediation method. The cost reduction of ADR is essential to make it a sustainable business. The major cost driver of ADR is the ADR spacecraft and its launcher, both of which are largely influenced by the spacecraft size. A trade-off study of the ADR spacecraft system has been conducted to minimize its size, resulting in the novel concept of a cost-effective ADR spacecraft system. The authors introduce the combination of electric propulsion (EP) and air drag utilization for the debris deorbit phase so as to lower the EP requirement, thereby leading to cost reduction of the ADR spacecraft. The minimum delta-V reentry method is also considered to reduce the propellant mass for final reentry. EP performance (e.g. thrust, specific impulse, mass) is modeled as a function of power consumption, and optimum EP performance has been explored to minimize overall mass of the ADR spacecraft, including its electrical power subsystem with this EP model. Given the minimum thrust limit to satisfy the mission duration requirement, a higher thrust-to-power ratio is preferable for debris removal electric propulsion. This paper discusses the optimal EP performance requirements in detail, and the results show that a roughly 500-kg ADR satellite using a 600-W-class Hall thruster can remove about three tons of a rocket body debris from an 800-km circular orbit within 600 days.

A robust controller for multi rotor UAVs

Unmanned aerial vehicles (UAVs) are safety-critical systems that often need to fly near buildings and over people under adverse wind conditions and hence require high manoeuvrability, accuracy, fast response abilities to ensure safety. Under extreme conditions, the dynamics of these systems are strongly nonlinear and are exposed to disturbances, which need a robust controller to keep the UAV and its environment safe. In this paper a novel robust nonlinear multi-rotor controller is introduced based on essential modifications of standard dynamic inversion control, which makes it insensitive to payload changes and also to large wind gusts. First a robust attitude controller is established, followed by lateral and vertical position control in a customary outer loop. The controllers take into account thrust limitations of the aircraft and theoretical proof is provided for robust performance. The control scheme is illustrated in simulation with a realistic nonlinear dynamical model of an aircraft that includes rotor dynamics and their speed limitations to show robustness. Lyapunov stability methods are used to prove the stability of the robust control system.

Optimal Concurrent Control for Space Manipulators Rendezvous and Capturing Targets Under Actuator Saturation

This article proposes a control method for space manipulators rendezvousing with and capturing a target, involving concurrent operation of an optimal and a coordinated controller. The optimal controller drives the whole system to rendezvous with the target, which saves onboard fuel and satisfies obstacle avoidance constraint and base spacecraft thrust limits. The optimal control problem is solved using calculus of variations method, with state inequality constraints transformed into extended dynamical subsystems and thrust limits formulated as saturation functions. The coordinated controller is designed based on the dynamic equations of the space manipulator on a noninertial frame attached to the system center of mass, which drives the end-effector to approach the target along a desired trajectory while making the base attitude follow a desired profile. It also generates proper reaction moments on base through manipulator motions to ensure controlling the base attitude when base attitude actuators reach their torque limits. The solution optimality of the optimal controller and the stability of the coordinated controller are demonstrated. Numerical simulations verify the performance of the proposed concurrent control method.

Neural network optimal control in astrodynamics: Application to the missed thrust problem

While high-efficiency propulsion techniques are enabling new mission concepts in deep space exploration, their limited thrust capabilities necessitate long thrusting arcs and make spacecraft more susceptible to missed thrust events. To correct for such mishaps, most spacecraft require updated trajectories that are relayed from Earth. While this solution is viable for spacecraft near Earth, in deep space, where one-way communication time is measured in hours, a delay in transmission may prolong the time of flight or result in a complete loss of mission. Such problems can be alleviated by increasing the spacecraft's onboard autonomy in guidance. This paper demonstrates how a computationally lightweight neural network can map the spacecraft's state to a near-optimal control action, autonomously guiding a spacecraft within different astrodynamic regimes and optimality criteria. The neural network is trained using supervised learning and datasets comprised of optimal state-action pairs, as determined through traditional direct and indirect methods. Additionally, the neural network-designed solutions retain optimality and time of flight corresponding to traditional trajectories. Finally, the same neural networks can autonomously correct for most missed thrust events encountered on long-duration low-thrust trajectories. The presented results provide a path for mitigating risks associated with the use of high-efficiency low-thrust propulsion techniques.

Nanosat Formation Flying Design for SNIPE Mission

This study designs and analyzes satellite formation flying concepts for the Small scale magNetospheric and Ionospheric Plasma Experiments (SNIPE) mission, that will observe the near-Earth space environment using four nanosats. To meet the requirements to achieve the scientific objectives of the SNIPE mission, three formation flying concepts are analyzed: a crossshape formation, a square-shape formation, and a cross-track formation. Of the three formation flying scenarios, the crosstrack formation scenario is selected as the final scenario for the SNIPE mission. The result of this study suggests a relative orbit control scenario for formation maintenance and reconfiguration, and the initial relative orbits of the four nanosats meeting the formation requirements and thrust limitations of the SNIPE mission. The formation flying scenario is validated by calculating the accumulated total thrust required for the four nanosats. If the cross-track formation scenario presented in this study is applied to the SNIPE mission, it is expected that the mission will be successfully accomplished.

Interpretation of gravity anomaly to delineate thrust faults locations at the northeastern part of India and its adjacent areas using source edge detection technique, tilt derivative and Cos(θ) analysis

The Northeast India and its adjacent areas converge among the three different plates, viz. Eurasia, India and Sunda plates. The tectonic interaction of Northeast India and underlying dynamics of the Himalayas as well as the Indo-Burma Ranges might cause the Assam Syntaxis. The area of study is located between latitude 23°–28°N and longitude 88°–96°E and situated in one of the most seismically active tectonic provinces in the world with seismic zone-V. This area had demonstrated several thrust faults activities and tectonic evident accomplishments during the recent past. The complicated geotectonic setups inspirits various smaller magnitude earthquakes, and the current seismicity shows seismic activities are still enduring in the Shillong Plateau, Arakan-Yoma fold belt, Bengal Basin, Naga Hills, Mikir Hills, Upper–Lower Brahmaputra Valley and Mismi Hills of Himalayan foothills. It is imperative to obtain wide-ranging learning tectonic configuration, thrust faults delineation for improved geoscientific study. Parts of the areas are extremely unreachable, and very limited thrust faults were marked by studying GIS map received from the various agencies and field geological study. During the past studies, most of the prominent lineaments/thrusts are marked; however, many active and hidden thrust faults are still unidentified. Seismic data can provide better information about the thrust faults locations, but due to small number of seismic data, the information is not adequate. In this paper, attempt has been made to study and reinterpret the available ground gravity data of northeastern parts of India for understanding thrust fault locations using various applications of gravity derivatives like analytical signal, horizontal gravity gradient, tilt derivative, horizontal tilt angle derivative and Cos(θ) analysis. Source edge detection technique has also been premeditated to categorize thrust fault locations. It is understandable that the low gravity is observed at Assam Valley which contributed sediment accumulations and higher gravity anomaly observed at Shillong Plateau and Bengal Basin containing denser formations. Bouguer gravity data is used after isostatic correction assuming Airy’s isostasy root depth model and first-order trend removal using least square technique. The derived thrust fault locations from the present study are superimposed with the existing thrust-fault locations for correlation. Some additional thrust faults are narrated which are not previously mapped. It is also suggested that Brahmaputra Thrust, Dauki Fault, Naga Thrust, Disang Thrust and Kopili Fault have key responsibility for high seismicity and tectonic movement causing upliftment and depression that encouraged some anticlockwise rotation in the area.

Mars Powered Descent Phase Guidance Design Based on Fixed-Time Stabilization Technique

This paper proposes a guidance scheme to achieve an autonomous precision landing on Mars and proposes a practical fixed-time stabilization theorem to analyze the robustness of the guidance. The proposed guidance is mainly based on the fixed-time stabilization method, and it can achieve the precision landing within a pre-defined time. This property enables the proposed guidance to outperform the finite-time stabilization technique which cannot handle uncertainties well and whose convergence time is dependent on initial states. Compared with the existing fixed-time stabilization theorem, the proposed practical fixed-time stabilization theorem can achieve a shorter convergence time and cope with unknown disturbances. When the Mars landing guidance is designed by this proposed theorem, the upper bound of the landing time and the maximum landing error subject to unknown disturbances can be calculated in advance. Theoretical proofs and Monte Carlo simulation results confirm the effectiveness of the proposed theorem and the proposed guidance. Furthermore, the efficacy of the proposed guidance with thrust limitations is also demonstrated by testing of 50 cases with a range of initial positions and velocities.

Fault Slip and Exhumation History of the Willard Thrust Sheet, Sevier Fold‐Thrust Belt, Utah: Relations to Wedge Propagation, Hinterland Uplift, and Foreland Basin Sedimentation

AbstractZircon (U‐Th)/He (ZHe) and zircon fission track thermochronometric data for 47 samples spanning the areally extensive Willard thrust sheet within the western part of the Sevier fold‐thrust belt record enhanced cooling and exhumation during major thrust slip spanning approximately 125–90 Ma. ZHe and zircon fission track age‐paleodepth patterns along structural transects and age‐distance relations along stratigraphic‐parallel traverses, combined with thermo‐kinematic modeling, constrain the fault slip history, with estimated slip rates of ~1 km/Myr from 125 to 105 Ma, increasing to ~3 km/Myr from 105 to 92 Ma, and then decreasing as major slip was transferred onto eastern thrusts. Exhumation was concentrated during motion up thrust ramps with estimated erosion rates of ~0.1 to 0.3 km/Myr. Local cooling ages of approximately 160–150 Ma may record a period of regional erosion, or alternatively an early phase of limited (<10 km) thrust slip. Propagation of the Sevier wedge front and major thrust slip during the late Early to mid‐Cretaceous were synchronous with increasing subsidence and deposition of thick synorogenic strata in the foreland, crustal thickening in the hinterland, growing igneous activity in the Sierran magmatic arc, and increasing plate convergence rates. Along‐strike, other parts of the Cordilleran retroarc fold‐thrust belt also experienced major shortening during the late Early to mid‐Cretaceous, following a period of earlier Cretaceous quiescence. Late Jurassic shortening was concentrated nearer the arc, and thus mostly in the hinterland at the latitude of northern Utah, related to width and location of the passive‐margin sedimentary wedge relative to the plate margin.

One-loop angularity distributions with recoil using Soft-Collinear Effective Theory

Angularities are event shapes whose sensitivity to the splitting angle of a collinear emission is controlled by a continuous parameter b, with −1 < b < ∞. When measured with respect to the thrust axis, this class of QCD observables includes thrust (b = 1) and jet broadening (b = 0), the former being insensitive to the recoil of soft against collinear radiation, while the latter being maximally sensitive to it. Presently available analytic results for angularity distributions with b ≠ 0 can be applied only close to the thrust limit since recoil effects have so far been neglected. As a first step to establish a comprehensive theoretical framework based on Soft-Collinear Effective Theory valid for all recoil-sensitive angularities, we compute for the first time angularity distributions at one-loop order in αs for all values of b taking into account recoil effects. In the differential cross section, these amount to novel sub-leading singular contributions and/or power corrections, where the former are characterized by fractional powers of the angularity and contribute appreciably close to the peak region, also for b ≳ 0.5. Our calculations are checked against various limits known in the literature and agree with the numerical output of the Event2 generator.

Aircraft Take-off and Landing Performance Calculation Method Based on Flight Simulation

In order to comply with the existing standard requirements or specifications, a new computational method for aircraft take-off and landing performance, which deals with the characteristics of the high precision of parameters in the process of take-off and landing based on flight simulation technique. The simulation model for a twin-engine normal layout aircraft is constructed in detail, including nonlinear motion equation, aerodynamic, engine, landing gear and dynamical mass model. According to performance calculation standard and pilot control specification for different take-off modes and landing stages, the simulation process for one engine inoperative(OEI) take-off, angle of attack(AoA) hold take-off, standard take-off, reject take-off and landing are designed, and corresponding performance computer software is developed to achieve the goal of accuracy as well as full parameters calculation. Compared with the existing methods, the computational complexity of this method is increased, the process is detailed, the parameters are increased, and more influencing factors can be analyzed quantitatively. Results show that OEI take-off distance is the longest, standard take-off is suitable for light aircraft, angle of attack hold take-off is appropriate in plateau or limited thrust. Landing distance is related to glide angle. Therefore, light aircraft landing can reduce the approach speed and increase the glide angle.

Dynamics and de-spin control of massive target by single tethered space tug

This paper investigates the dynamics and de-spin control of a massive target by a single tethered space tug in the post-capture phase. The dynamic model of the tethered system is derived and simplified to a dimensionless form. Further, a decoupled PD controller is proposed, and the local stability of the controller is analyzed by linearization technique. Parametric studies of the dynamics and de-spin control of a massive target are conducted to characterize the dynamic process of de-spin with the proposed control law. It is shown that the massive target can be de-span by a single and small space tug with limited thrust within finite time. The thrust tangent with the tether de-spins the target while the thrust normal to the tether prevents the tether from winding up the target. The tether length has a positive contribution to the de-spin of a target. The longer tether leads to a faster de-spin process.

Pulsed Optimal Spacecraft Orbit Reorientation by Means of Reactive Thrust Orthogonal to the Osculating Orbit. II

The first part of the article provides an overview of the work on the differential equations of the spacecraft (SC) orbit orientation and the problem of optimal reorientation of a spacecraft orbit in an inertial coordinate system by means of reactive acceleration orthogonal to the osculating plane of the spacecraft. The theory of solving the problem of the optimal reorientation of the orbit of the spacecraft using the quaternionic differential equation for the orientation of the orbital coordinate system in a non-linear continuous formulation (using limited (small) thrust) is presented. As a minimized quality functional, a combined functional is used equal to the weighted sum of the reorientation time and thrust impulse (characteristic speed) during the reorientation of the orbit of the spacecraft (special cases of this functional are the speed response case and the characteristic speed minimization separately).

A Performance Analysis Method of High Speed and Small Diameter Propeller

Recently, water-air drone has shown widespread applications in various fields, such as military and live entertainment. As a direct source of power, it is necessary to study and calculate the performances of water-air propeller. The underwater propeller has the maximum diameter limit and the air propeller has the minimum thrust limit. So, high speed and small diameter are the characteristics of water-air propeller. In this paper, the performance in air of a water-air propeller has been calculated using programs, which were based on the strip theory. The effect of rotate speed on thrust, torque and power-to-thrust ratio has been carefully investigated. Meanwhile, these results have been verified by physical experiments.

Use of Programs for Two-Parameter Multiple Impulse Correction of Altitude and Inclination of Circular Orbits

The paper describes the method for the development of programs for two-parameter multiple impulse correction of altitude and inclination of circular orbits. The considered corrections feature the simultaneous adjustment of altitude and inclination of the orbit under limited thrust of the propulsion system of the spacecraft. Low thrust-to-weight ratio of the spacecraft leads to the need for a correction program consisting of several burns of the propulsion system. Thrust vector orientation, burn time, and its operation duration are determined as propulsion system parameters.

Reactive Trajectory Generation for Multiple Vehicles in Unknown Environments With Wind Disturbances

Unmanned aerial vehicle use continues to increase, including operating beyond line of sight in unknown environments where the vehicle must autonomously generate a trajectory to safely navigate. In this article, we develop a trajectory generation algorithm for vehicles with second-order dynamics in unknown environments with bounded wind disturbances in which the vehicle only relies on its on-board distance sensors and communication with other vehicles to navigate. The proposed algorithm generates smooth trajectories and can be used with high-level planners and low-level motion controllers. The algorithm computes a maximum safe cruise velocity for the vehicle in the environment and guarantees that the trajectory does not violate the vehicle's thrust limitation, sensor constraints, or user-defined clearance radius around other vehicles and obstacles. Additionally, the trajectories are guaranteed to reach a stationary goal position in finite time given a finite number of bounded obstacles. Simulation results demonstrate the algorithm properties through two scenarios. First, a quadrotor navigating through a moving obstacle field to a goal position and, second, multiple quadrotors navigating into a building to different goal positions.

Use of Dynamic Scaling for Trajectory Planning of Floating Pedestal and Manipulator System in a Microgravity Environment

In this paper, motion planning and coordination is investigated for a space robot composed of a floating pedestal and manipulator. In some cases, such as a manipulator grasping a higher quality target, the dynamic coupling can occur leading to under-actuation of the floating pedestal (that is, the required control force of the pedestal exceeds the thrust limit). As a result, the desired operation may not be achieved due to large control error. Therefore, we propose an innovative planning method, termed dynamic scaling planning method, to avoid pedestal under-actuation and guarantee accuracy of manipulator operations. Furthermore, to validate the proposed method, an experimental model of a space robot operating in a magnetic-liquid hybrid suspension microgravity simulation environment was developed. Results of the experimental simulations demonstrate that the proposed method can effectively avoid under-actuation of the pedestal. Moreover, the end-effector of the manipulator follows a desired path to successfully reach its target location.

The use of three-impulse transfer to insert the spacecraft into the high Moon Artificial Satellite orbits

The problem of the optimal spacecraft’s insertion from the Earth into the high circular polar Moon Artificial Satellite’s orbit (MAS) with a radius of 4000–8000 km has been investigated. A comparison of single- and three-impulse insertion schemes has been performed. The analysis was made taking into account the disturbances from the lunar gravity field harmonics and the gravity fields of the Earth and the Sun, as well as the engine’s limited thrust. It has been shown that the three-impulse transfer from the initial selenocentric hyperbola of the approach into the considered final high MAS orbit is noticeably better with respect to the final mass than the ordinary single-impulse deceleration. The control parameters that implement this maneuver and provide nearly the same energy expenses as in the Keplerian case have been presented. It was found that, in contrast to the Keplerian case, in the considered case of the real gravity field, there is the optimal maximum distance of the maneuver. Recently, the Moon exploration problem became actual again.

Minimum time multiple-burn optimization of an upper stage with a finite thrust for satellite injection into geostationary orbit

In this paper, the problem of minimum time multiple-burn optimization of an upper stage with a limited thrust, and engine restart capability for satellite injection into geostationary orbit are considered. The goals are to find thrust vector angle, times of the engine firings, and optimal duration of active phases of the upper stage to minimize fuel consumption and meet the desired boundary conditions. Various flight sequences with multiple burns, from two burns up to six burns, are considered. Also, the optimal trajectory for each sequence is derived. To solve the multi-point boundary value problem, an improved indirect shooting method with high performance is presented and used for an optimal solution. All in all, this novel method presented for multi-burn problem, not only with a very good accuracy, but also with a very fast convergence to the desired end conditions.

A new approach to autonomous rendezvous for spacecraft with limited impulsive thrust: Based on switching control strategy

This paper proposes a novel impulsive control approach for autonomous rendezvous of spacecraft with limited thrust. Based on the dynamic model established by Clohessy–Wiltshire (C–W) equations, the rendezvous process is regarded as impulse phase and free motion phase which are considered as closed-loop system and open-loop system respectively. The durations of the two phases are scaled by sampling period. Then, based on Lyapunov theory, two virtual energy functions of the subsystems are introduced, and the rendezvous problem is regarded as an asymptotic stabilization problem of the switching system. Meanwhile, the thrust limitation is also taken into consideration. The impulsive closed-loop control problem is finally transformed into a convex optimization problem and the calculation steps are proposed. With the designed controller, the impulsive thrust during the rendezvous is calculated at the impulse instants. Furthermore, the calculated thrust is kept below a given bound, which can also be minimized according to specific requirements. The effectiveness of the proposed approach is illustrated by simulation examples.

Analysis of Force-Capping for Large Wind Turbine Rotors

A standard design procedure is to optimise wind turbine rotors for maximum power coefficient within a specified range of wind speeds, up to the rated wind speed. The downwind thrust force upon the rotor is greatest at this design condition. Reacting large forces using cantilever beams necessitates high material costs for the blades, tower and foundations. This paper shows that limiting the maximum allowable downwind thrust can lower the capital expenditure (CAPEX) of a wind farm per MW installed. This saving is made at the expense of sacrificing power coefficient over the small range of wind speeds at which the thrust limit is enforced. We describe this concept as ‘force-capping’. For small reductions in the maximum allowable force, the effect on annual energy production (AEP) is minor. This paper outlines a parametric analysis procedure to evaluate how reducing the material content of a wind turbine affects the energy production. Dynamic forces and fatigue life are considered. For small reductions in turbine material, we show that the change in CAPEX is first order, whereas the loss of AEP is second order. We therefore conclude that force-capping must be beneficial in every case to at least some extent. Our analysis shows that the cost of energy can reduce by approximately 0.2% through imposing a force-cap of ∼90% upon the maximum thrust force for a 5MW machine. The case for force-capping strengthens further as the cost of borrowing and cost of material rises.