Attitude Control Thrusters Research Articles

Constrained Station Relocation in Geostationary Equatorial Orbit Using a Legendre Pseudospectral Method

An algorithm for numerical approximation is formulated using a Legendre pseudospectral method to solve the problem of constrained station relocation, changing current position to another longitude, in geostationaryequatorial orbit. The problem was stated as a Bolza form in optimal control theory to calculate the control. A collision avoidance term, which is included in the objective function as an integral form and serves as a soft constraint, is considered during transfer by specifying a maximum or minimum radius, depending on whether the transfer is in the east or west direction. Several east-and-west-direction station relocation maneuvers are simulated; the states and control behavior are obtained and all the variables are found to be within feasible ranges. This method would be possible to implement using very-low-thrust engines, in particular, by employing attitude control thrusters.

70 N급 하이드라진 추력기의 촉매대 형상(L/D) 최적화 연구

우주비행체의 주 엔진 및 자세제어용 추력기로 적용 가능한 단일액체추진제 하이드라진 추력기의 촉매대 최적설계형상 도출을 위한 지상연소시험을 수행하였다. 추력, 비추력, 연소실 압력, 특성속도, 그리고 하이드라진 분해반응률 등을 추력실 길이변화에 따라 고찰하고, 추진제 공급압력과 성능변수가 갖는 상관관계를 제시한다. 성능평가결과, 시험조건 내에서의 촉매대 길이증가는 추력기의 성능저하를 야기하고, 공급압력 변이에 따른 추진제 소모 효율성 또한 떨어뜨리는 것이 확인되었다. A ground hot-firing test was conducted to take out the optimal design configurations for the catalyst bed of liquid-monopropellant hydrazine thruster which could be used for primary engine or attitude control thruster of space vehicles. Performance characteristics with the variation of thrust-chamber length are investigated in terms of thrust, specific impulse, chamber pressure, characteristic velocity, and hydrazine decomposition rate. Additionally, the correlations between propellant-supply pressure and performance parameters are given. As results, increase of catalyst-bed length leads to performance degradation in this test condition, and also decreases propellant consumption efficiency with the supply pressure variation.

Research on vacuum plume and its effects

In vacuum environment, the exhaust flow of attitude control thrusters would expand freely and produce the plume, which possibly causes undesirable contamination, aerodynamic force and heating effects to the spacecraft. Plume work station (PWS) is developed by Beihang University (BUAA) for numerically simulating the vacuum plume and its effects. An approach which combines the direct simulation Monte Carlo (DSMC) method and difference solution of Navier–Stokes (N–S) equations is applied. The internal flows in nozzles are simulated by solving the NS equations. The flow parameters at nozzle exit are used as the inlet boundary condition for the DSMC calculation. Experimental studies are carried out in a supersonic low density wind tunnel which could simulate the 60–80 km altitude environment to investigate the plume and its effects. To demonstrate the capability of PWS, numerical simulations are performed for the vacuum plume of several typical attitude control thrusters. The research results are of great help for the engineering design.

System Dynamic Characteristic Simulation of Spacecraft Propulsion System Based on AMESim

In order to investigate the dynamic characteristic of the spacecraft propulsion system, numerical simulation was carried out based on the modular and extensible simulation model UPSSim. The model accurately predicts the global behavior of the spacecraft propulsion system in the priming and shutdown process, and also correctly describes the transient behavior of the subsystems. Due to the distributaries of the attitude control thruster supply pipelines, the propellant flow fluctuation attenuates rapidly in the priming process and the waterhammer peak pressure of main engine supply pipelines decreases significantly. The waterhammer waves of the main supply pipelines propagate to the attitude control thruster supply pipelines and dissipate gradually. The farther from the main engine, the later the pressure fluctuates.

Tradeoff Analysis of Attitude-Control Slew Algorithms for Prolate Spinner

There have been a number of slew algorithms in literature such as half-cone and rhumb line that allow attitude control of a prolate spinning spacecraft using only a single attitude-control thruster. Recently, several novel slew algorithms have been further introduced to the field: extended half-cone, dual cone, and spin sync, which offer a number of advantages to the existing algorithms. The ultimate output of this paper is a slew algorithm selection tradeoff, which is performed in the context of a case study of theMoonLightweight Interior andTelecomExperiment interplanetary penetrator mission. To support this final quantitative tradeoff, the following analyses are presented beforehand: 1) a summary of the state-of-the-art slew algorithms, 2) a thorough generic parametric analysis of relevant characteristic performances (slew accuracy, slew duration, energy consumption, and torque capacity) as influenced by relevant mission variables, and 3) a comprehensive slew algorithm selection tradeoff methodology. As the parametric analyses and tradeoffmethodology are generic in nature, this work could be reused to support similar tradeoffs for other mission scenarios; this represents a major added value of this study. The numerical results presented are validated using a newly developed software simulator called the Attitude Control Simulator for Spinning Spacecraft with Single Thruster.

Analysis of state-of-the-art single-thruster attitude control techniques for spinning penetrator

The attitude dynamics and manoeuvre survey in this paper is performed for a mission scenario involving a penetrator-type spacecraft: an axisymmetric prolate spacecraft spinning around its minor axis of inertia performing a 90 spin axis reorientation manoeuvre. In contrast to most existing spacecraft only one attitude control thruster is available, providing a control torque perpendicular to the spin axis. Having only one attitude thruster on a spinning spacecraft could be preferred for spacecraft simplicity (lower mass, lower power consumption etc), or it could be imposed in the context of redundancy/ contingency operations. This constraint does yield restrictions on the thruster timings, depending on the ratio of minor to major moments of inertia among other parameters. The Japanese Lunar-A penetrator spacecraft proposal is a good example of such a single-thruster spin-stabilised prolate spacecraft. The attitude dynamics of a spinning rigid body are first investigated analytically, then expanded for the specific case of a prolate and axisymmetric rigid body and finally a cursory exploration of non-rigid body dynamics is made. Next two well-known techniques for manoeuvring a spin-stabilised spacecraft, the Half-cone/Multiple Half- cone and the Rhumb line slew, are compared with two new techniques, the Sector-Arc Slew developed by Astrium Satellites and the Dual-cone developed at Surrey Space Centre. Each technique is introduced and characterised by means of simulation results and illustrations based on the penetrator mission scenario and a brief robustness analysis is performed against errors in moments of inertia and spin rate. Also, the relative benefits of each slew algorithm are discussed in terms of slew accuracy, energy (propellant) efficiency and time efficiency. For example, a sequence of half-cone manoeuvres (a Multi-half-cone manoeuvre) tends to be more energy-efficient than one half-cone for the same final slew angle, but more time-consuming. As another example, the new Sector-Arc Slew and Dual-cone techniques are designed to overcome a specific restriction on attainable slew angle that is associated with the half-cone manoeuvre, giving one additional degree of freedom for designers to fine-tune. © 2012 Elsevier Ltd.

A Kinematic Approach to Determining the Optimal Actuator Sensor Architecture for Space Robots

Autonomous space robots will be required for such future missions as the construction of large space structures and repairing disabled satellites. These robots will need to be precisely controlled. However, factors such as manipulator joint/actuator friction and spacecraft attitude control thruster inaccuracies can substantially degrade control system performance. Sensor-based control algorithms can be used to mitigate the effects of actuator error, but sensors can add substantially to a space system’s weight, complexity, and cost, and reduce its reliability. Here, a method is presented to determine the sensor architecture that uses the minimum number of sensors that can simultaneously compensate for errors and disturbance in a space robot’s manipulator joint actuators, spacecraft thrusters, and reaction wheels. The placement and minimal number of sensors is determined by analytically structuring the system into “canonical chains” that consist of the manipulator links and spacecraft with force/torque sensors placed between the space robot’s spacecraft and its manipulators. These chains are combined to determine the number of sensors needed for the entire system. Examples of one- and two-manipulator space robots are studied and the results are validated by simulation.

Orbit Maneuver Compensation of KAGUYA for Its Safe and Accurate Lunar Transfer

Reported in this paper are the results of the orbit maneuver compensation in KAGUYA's Lunar transfer. Because of the uncoupled allocation of the attitude control thrusters, extra velocity increment (δv ) is induced whenever KAGUYA performs an orbit maneuver. Since the observed level of δv was unacceptable range from the point of maneuver accuracy requirement, it was compensated by means of deducting estimated δv from the orbit maneuver command. The δv estimation model was updated step-by-step during the Lunar transfer, which leaded to significant improvement of the orbit maneuver accuracy and resulted in the omission of the last trajectory correction maneuver. The method of the compensation and its results are introduced in detail.

Reliable fault diagnosis scheme for a spacecraft attitude control system

This paper presents a scheme for fault detection and isolation (FDI) of on-board gyroscope sensors and thrusters for spacecraft attitude control, based on the example of the Mars Express (MEX) satellite. The main contribution of the paper is related to the design and the optimization of an FDI procedure based on robust observers or filters, used as estimators, which generate the FDI residual signals. When organized into an estimator bank, excellent fault isolation properties are achieved upon suitable design. The residual evaluation relies on decision logic, whose thresholds are properly selected and specified. The FDI strategy is applied to the non-linear simulation of the MEX system, and the FDI performance is evaluated subject to disturbance signals, model uncertainty, and measurement noise processes. The robustness and reliability properties of the robust residual generators are investigated and verified in simulation by selecting suitable performance criteria together with Monte Carlo analysis. The results obtained highlight a good trade-off between solution complexity and achieved performances. Comparisons with the existing fault diagnosis algorithms implemented on board the MEX spacecraft are finally reported. The proposed FDI design methodology constitutes a reliability approach for real application of FDI in future spacecraft.

Thermal Analysis for Propellant Stream in Thruster'sInjection Tube During Start Process

According to the leading design features of the monopropellant attitude-control thruster and the cryogenic condition in space, the coupling heat-transfer physical model and a set of mathematical equations are put forward to calculate the temperature variation of the injection tube and three other components of the thruster, which are the aggregate organ, the bracket, and the injection plate. The wall temperature of the injection tube is taken as the thermal boundary conditions for the propellant stream temperature drop calculation. The temperature calculation of the propellant stream flowing inside the injection tube during the start process in a typical 10 N thruster has been performed, and two simulation experiments are carried out on the same type thruster in the vacuum cryogenic chamber. The results of the calculation agree well with the experiments. The validated calculation model can be used to judge whether the propellant stream freezes or not in the injection tube during the start process after the thruster has been exposed to the cryogenic space for a long time. In addition, four factors that affect the front temperature of the propellant stream at the injection tube exit during the propellant filled process are studied, and according to the calculation results, three correlation relationships are deduced, respectively.

N-Nitrosodimethylamine Release from Fuel Oxidizer Reaction Product Contamined Extravehicular Activity Suits

Before extravehicular activity (EVA) on the Russian segment (RS) of the International Space Station (ISS), the docking compartment (DC1) must be depressurized, because it is used as an airlock. It is preferred to use the U.S. control moment gyros (CMGs) instead of attitude-control thruster firings to compensate for disturbances and to maintain the ISS vehicle attitude. However, when the DC1 is depressurized, the CMGs’ margin of momentum is insufficient to compensate for the disturbance and the service module (SM) attitude-control thrusters’ need to fire to desaturate the CMGs. The SM roll-control thruster firings induce fuel‐oxidizer reaction products (FORP) contamination on the adjacent SM surfaces around the thrusters. One of the components present in FORP is the potent carcinogen N-nitrosodimethylamine (NDMA). Because the EVA crewmembers often enter the area surrounding the thrusters for tasks on the aft end of the SM and when translating to other areas of the RS, the presence of FORP contamination is a concern. FORP contamination of the SM surfaces is discussed, along with the potential release of NDMA in a humid environment from crew EVA suits, whether they happen to be contaminated with FORP, the toxicological risk associated with the NDMA release, and the implementation of flight rules to mitigate the hazard.

A MONTE CARLO ANALYSIS AND DESIGN FOR FDI OF A SATELLITE ATTITUDE CONTROL SYSTEM

This paper deals with the reliability and performance evaluation of two methods for FDI of on-board sensors and thrusters for attitude control of the Mars Express (MEX) orbiter spacecraft. The study makes use of a detailed nonlinear simulation of the MEX, and the FDI performance is evaluated subject to disturbance signals, model uncertainty and measurement noise processes. The two methods are based on robust dynamic observers or filters, used as estimators, which generate the FDI residual signals. When organised into an estimator bank excellent fault isolation properties are achieved upon suitable design. Carefully selected performance criteria indices are used together with Monte Carlo robustness tuning and performance evaluation. These constitute the FDI design methodology realising a reliability approach for real application of FDI in future spacecraft.

Simulations of Cold-Gas Nozzle and Plume Flows and Flight Data Comparisons

The nozzle and plume  ows of small cold-gas attitude control thrusters, plume interactions with spacecraft surfaces, and the induced pressure environment are investigated numerically. The motivation for this study originates from pressure measurements that exhibited nonperiodic pulses during the Ž rings of small cold-gas thrusters onboard a suborbital spacecraft. Pitch, yaw, and roll cold-gas thrusters were located on the 0.56-m-diam base of the spacecraft. The conical spacecraft  ew at altitudes between 670 and 1200 km and carried inside a pressure sensor connected to the side surface with a tube. Predictions of the pressure inside the sensor chamber are obtained using a semi-analyticalmodel with inputs from coupled continuumand kinetic simulations.The nozzle and plume  ows for each thruster are simulated using a three-dimensional Navier–Stokes solver until breakdown. FlowŽ eld properties inside the breakdown surface are used as inputs to the direct simulationMonte Carlo calculations in a domain that includes the spacecraft geometry. FlowŽ eld properties at the entrance of the sensor tube are used as inputs to an analyticalmodel to obtain the pressure inside the sensor chamber. Simulationsshow plume expansion, re ection off the spacecraft surfaces, and back ow. Pressure predictions for the pitch and yaw thruster plumes that reach the sensor after expanding on the spacecraft base are in very good agreement with measurements. Pressure induced by the roll thrusters is shown to be very sensitive to their radial position at the Environmental Monitor Package base and decreases with decreasing radial distance. Pressure overprediction of the roll thrusters is attributed to possible difference between the simulated and actual radial position.

On-orbit thruster calibration

Significant errors in force and moment output as well as large control coupling forces can result if the attitude and translation control thrusters of a spacecraft are not calibrated properly. Thruster calibration is also important for estimating fuel usage so that spacecraft life can be accurately predicted. A general procedure to accurately determine the true relationship between the commanded and the actual force output of a set of thrusters is presented. This relationship is determined from data generated by a spacecraft while it is in orbit. The calibration is based on the Kalman filter estimator and is verified by a digital computer simulation of the Gravity Probe B spacecraft dynamics. Gravity Probe-B is an Earth orbiting gravitational physics experiment, which will test several aspects of Einstein's general relativity theory. A set of 18 proportional thrusters generate attitude control moments around three axes and translation control forces along three axes for full six-degree-of-freedom control. A total of 108 parameters are needed to characterize both the force and moment outputs of all 18 thrusters. The ability to calibrate these parameters to an accuracy of better than 1 % rms is demonstrated.

Bipropellant propulsion performance and propellant-remaining prediction for INSAT-1B

The INSAT-1B communications and weather satellite was launched from the Space Shuttle Challenger on August 31, 1983. The satellite has been operating successfully in orbit for more than 5 years and 10 months. It provides extremely valuable flight data to characterize the performance of a bipropellant propulsion subsystem for both the transfer orbit and on-orbit operations. Analytical modes have been developed and calibrated to determine the performance of the 490-N apogee thruster and the 22-N attitude and orbit control thrusters, the plume-impingements effect on the reduction of thruster velocity increment efficiency, and the propellant remaining. These flight-proven analytical models have demonstrated the ability to predict and simulate in-flight performance with a high degree of accuracy. The integrated bipropellant propulsion subsystem and the flight data will be briefly described. Subsystem performance is illustrated by comparing the flight data to the analytical models.

A technique for the fine longitude control of a geostationary satellite as demonstrated on insat‐1B

AbstractThe possibilities of finer longitude control of a geostationary satellite are studied with a view to providing insight into the problem of co‐location of satellites. It is shown that for a three‐axis stabilized satellite such as INSAT‐1B, placed close to an equilibrium longitude, the propellant expenditure for satellite attitude control is considerably larger than that required for longitude drift control. An experiment was conducted on INSAT‐1B in which the forces arising from the attitude control thrusters are optimally used to counteract the force due to the solar radiation pressure thereby reducing the perturbations on the orbital eccentricity vector. This is accomplished by suitably switching between the redundant attitude control thrusters located in the positive and negative roll faces of the satellite. The extent to which the effects of the solar radiation pressure are nullified is quantitatively estimated from the observed behaviour of the eccentricity vector. Without any recurrent propellant expenditure it is shown that it is possible to restrict the longitudinal excursions to nearly half of the normally allocated ±0·1° longitude band.

Vehicle‐environment interaction workshop

On February 20–21, 1988, the Vehicle‐Environment Interaction Workshop was held at the Applied Physics Laboratory, Johns Hopkins University, Laurel, Md., to evaluate the current understanding of the physical and chemical processes that occur when vehicles in low Earth orbit (LEO) interact with the ambient atmosphere, the unique characteristic of this environment being 8‐km/s impact velocity of the incoming atoms and molecules. Most of the papers on the first day were devoted to the phenomenology and interpretation of space vehicle glows, whereas on the second day, laboratory investigations and discussions of charged particle interactions were presented.The unambiguous recognition of the phenomenon of spacecraft glows occurred when the astronauts on shuttle flight STS‐3, in March 1982, visually observed an orange glow on the shuttle surfaces visible from the flight deck [Banks et al., 1983]. Photographs showed that the glow was strongest on surfaces facing into the incoming atmosphere (the ram direction) and that it decayed with increasing distance from the surface, with a characteristic length on the order of 20 cm. The glow intensity was seen to be strongly enhanced when the rear attitude control thrusters were fired, when both surface‐related and homogenous glows were observed.

Attitude control thruster plume flow modeling and experiments

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

The shuttle environment: Effects of thruster firings on gas density and composition in the payload bay

In order to study the gaseous environment of the STS orbiter a mass spectometer was operated aboard the SPAS‐01 subsatellite on the shuttle missions STS 7 and 41‐B. Data are reported on the impact of the operation of attitude control thrusters on the atmosphere in the payload bay. Firings of aft mounted thrusters have been observed to increase the total number density inside the payload bay to more than 1018 m−3 (equivalent to more than 7×10−5 mbar), whereas the operation of most of the forward mounted thrusters yields hardly a detectable signal. Nitric oxide is a major gaseous constituent inside the payload bay during the operation of aft mounted thrusters. Water desorbing from surfaces leads to a slowly decaying total number density following engine shutdown.

Empirical simulation model for hydrazine attitude control thrusters

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.