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InSight Attitude Control System Thruster Characterization and Calibration for Successful Navigation to Mars
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TRMM Re-Entry Planning: Attitude Determination and Control During Thruster Modes
The Tropical Rainfall Measuring Mission (TRMM) spacecraft has been undergoing design for a controlled re-entry to Earth. During simulation of the re-entry plan, there was evidence of errors in the attitude determination algorithms during thruster modes. These errors affected the bum efficiency, and thus planning, during re-entry. During thruster modes, the spacecraft attitude is controlled off of integrated Gyro Error Angles that were designed to closely follow the nominal spacecraft pointing frame (Tip Frame). These angles, however, were not exactly mapped to the Tip Frame from the Body Frame. Additionally, in the initial formulation of the thruster mode attitude determination algorithms, several assumptions and approximations were made to conserve processor speed. These errors became noticeable and significant when simulating bums of much longer duration (-10 times) than had been produced in flight. A solution is proposed that uses attitude determination information from a propagated extended Kalman filter that already exists in the TRMM thruster modes. This attitude information is then used to rotate the Gyro Error Angles into the Tip Frame. An error analysis is presented that compares the two formulations. The new algorithm is tested using the TRMM High-Fidelity Simulator and verified with the TRMM Software Testing and Training Facility. Simulation results for both configurations are also presented.
Minimum impulse technologies for precision attitude control: Minimum Impulse Thruster (MIT) and Hydrazine milliNewton Thruster (HmNT)
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Lunar Reconnaissance Orbiter (LRO) Rapid Thermal Design Development
The Lunar Reconnaissance Orbiter (LRO) project had a rapid development schedule starting with project conception in spring of 2004, instrument and launch vehicle selection late in 2005 and then launch in early 2009. The lunar thermal environment is one of the harshest in our solar system with the heavy infrared loading of the moon due to low albedo, lack of lunar atmosphere, and low effective regolith conduction. This set of constraints required a thermal design which maximized performance (minimized radiator area and cold control heater power) and minimized thermal hardware build at the orbiter level (blanketing, and heater service). The orbiter design located most of the avionics on an isothermalized heat pipe panel called the IsoThermal Panel (ITP). The ITP was coupled by dual bore heat pipes to an Optical Solar Reflector (OSR) covered heat pipe radiator. By coupling all of the avionics to one system, the hardware was simplified. The seven instruments were mainly heritage instruments which resulted in their desired radiators being located by their heritage design. This minimized instrument redesigns and therefore allowed them to be delivered earlier, though it resulted in a more complex orbiter level blanket and heater service design. Three of the instruments were mounted on a tight pointing M55J optical bench that needed to be covered in heaters to maintain pointing. Two were mounted to spacecraft controlled radiators. One was mounted to the ITP Dual Bores. The last was mounted directly to the bus structure on the moon facing panel. The propulsion system utilized four-20 pound insertion thrusters and eight-5 pound attitude control thrusters (ACS) in addition to 1000 kg of fuel in two large tanks. The propulsion system had a heater cylinder and a heated mounting deck for the insertion thrusters which coupled most of the propulsion design together simplifying the heater design. The High Gain Antenna System (HGAS) and Solar Array System (SAS) used dual axis actuator gimbal systems. HGAS required additional boom heaters to cool the approximately 10 W of RF losses thru the rotary joints and wave guides from the 40 W Ka system. By design this module needed a fair amount of heater, blanketing, and radiator complexity. The SAS system required a separate cable wrap radiator to help cool the Solar Array harness which dissipated 30 W thru the actuators and cable wraps. This module also was complex.
Space Technology: Game Changing Development Deep Space Engine (DSE) 100 lbf and 5 lbf Thruster Development and Qualification
Science mission studies require spacecraft propulsion systems that are high-performance, lightweight, and compact. Highly matured technology and low-cost, short development time of the propulsion system are also very desirable. The Deep Space Engine (DSE) 100-lbf thruster is being developed to meet these needs. The overall goal of this game changing technology project is to qualify the DSE thrusters along with 5-lbf attitude control thrusters for space flight and for inclusion in science and exploration missions. The aim is to perform qualification tests representative of mission duty cycles. Most exploration missions are constrained by mass, power and cost. As major propulsion components, thrusters are identified as high-risk, long-lead development items. NASA spacecraft primarily rely on 1960s' heritage in-space thruster designs and opportunities exist for reducing size, weight, power, and cost through the utilization of modern materials and advanced manufacturing techniques. Advancements in MON-25/MMH hypergolic bipropellant thrusters represent a promising avenue for addressing these deficiencies with tremendous mission enhancing benefits. DSE is much lighter and costs less than currently available thrusters in comparable thrust classes. Because MON-25 propellants operate at lower temperatures, less power is needed for propellant conditioning for in-space propulsion applications, especially long duration and/or deep-space missions. Reduced power results in reduced mass for batteries and solar panels. DSE is capable of operating at a wide propellant temperature range (between -22 F and 122 F) while a similar existing thruster operates between 45 F and 70 F. Such a capability offers robust propulsion operation as well as flexibility in design. NASA's Marshall Space Flight Center evaluated available operational Missile Defense Agency heritage thrusters suitable for the science and lunar lander propulsion systems.
Spacecraft contamination prediction and testing techniques
Techniques used in the prediction of spacecraft contamination for the Galileo Jupiter Orbiter and in the determination of the effects of such contamination are presented. Following a quick-look assessment of the contributions of ground-based initial contaminant loading, launch vehicle interface effects, vacuum-exposed outgassing deposition and attitude control thruster impingement and venting to the spacecraft contamination burden, the evaluations centered on the effects of the attitude control thruster on the scan platform optics, including calculations of thruster flowfields and a high-fidelity computer simulation of contaminant distribution. The evaluations revealed a considerable problem with thruster contamination, which could be solved by the use of a thrust shield and the avoidance of thruster operation at certain scan platform orientations. The effects of the various possible contaminants on spacecraft thermal and optical system performances were also investigated in studies of the optical transmittance of deposited monomethyl hydrazine nitrate, vacuum optical degradation due to contaminant outgassing and re-emission outgassing, and an operational satellite contaminant monitor on the NOAA-C satellite. It is concluded that with a good evaluation and testing program, contamination control may become a necessary portion of system design procedures, and recommendations for the implementation of various practices and tests to minimize contamination effects are presented.
Fuzzy based attitude controller for flexible spacecraft with on/off thrusters
A fuzzy-based attitude controller is designed for attitude control of a generic spacecraft with on/off thrusters. The controller is comprised of packages of rules dedicated to addressing different objectives (e.g., disturbance rejection, low fuel consumption, avoiding the excitation of flexible appendages, etc.). These rule packages can be inserted or removed depending on the requirements of the particular spacecraft and are parameterized based on vehicle parameters such as inertia or operational parameters such as the maneuvering rate. Individual rule packages can be 'weighted' relative to each other to emphasize the importance of one objective relative to another. Finally, the fuzzy controller and rule packages are demonstrated using the high-fidelity Space Shuttle Interactive On-Orbit Simulator (IOS) while performing typical on-orbit operations and are subsequently compared with the existing shuttle flight control system performance.
Fuzzy based attitude controller for flexible spacecraft with on/off thrusters
A fuzzy-based attitude controller is designed for attitude control of a generic spacecraft with on/off thrusters. The controller is comprised of packages of rules dedicated to addressing different objectives (e.g., disturbance rejection, low fuel consumption, avoiding the excitation of flexible appendages, etc.). These rule packages can be inserted or removed depending on the requirements of the particular spacecraft and are parameterized based on vehicle parameters such as inertia or operational parameters such as the maneuvering rate. Individual rule packages can be 'weighted' relative to each other to emphasize the importance of one objective relative to another. Finally, the fuzzy controller and rule packages are demonstrated using the high-fidelity Space Shuttle Interactive On-Orbit Simulator (IOS) while performing typical on-orbit operations and are subsequently compared with the existing shuttle flight control system performance.
Voyager Uranus encounter 0.2lbf T/VA short pulse test report
The attitude control thrusters on the Voyager spacecraft were tested for operation at electrical pulse widths of less than the current 10-millisecond minimum to reduce impulse bit and, therefore, reduce image smear of pictures taken during the Uranus encounter. Thrusters with the identical configuration of the units on the spacecraft were fired in an altitude chamber to characterize impulse bit and impulse bit variations as a function of electrical pulse widths and to determine if the short pulses decreased thruster life. Pulse widths of 4.0 milliseconds provide approximately 45 percent of the impulse provided by a 10-ms pulse, and thruster-to-thruster and pulse-to-pulse variation is approximately plus or minus 10 percent. Pulse widths shorter than 4 ms showed wide variation, and no pulse was obtained at 3 ms. Three thrusters were each subjected to 75,000 short pulses of 4 ms or less without performance degradation. A fourth thruster exhibited partial flow blockage after 13,000 short pulses, but this was attributed to prevous test history and not short pulse exposure. The Voyager attitude control thrusters should be considered flight qualified for short pulse operation at pulse widths of 4.0 ms or more.
ISS Contingency Attitude Control Recovery Method for Loss of Automatic Thruster Control
In this paper, the attitude control issues associated with International Space Station (ISS) loss of automatic thruster control capability are discussed and methods for attitude control recovery are presented. This scenario was experienced recently during Shuttle mission STS-117 and ISS Stage 13A in June 2007 when the Russian GN&C computers, which command the ISS thrusters, failed. Without automatic propulsive attitude control, the ISS would not be able to regain attitude control after the Orbiter undocked. The core issues associated with recovering long-term attitude control using CMGs are described as well as the systems engineering analysis to identify recovery options. It is shown that the recovery method can be separated into a procedure for rate damping to a safe harbor gravity gradient stable orientation and a capability to maneuver the vehicle to the necessary initial conditions for long term attitude hold. A manual control option using Soyuz and Progress vehicle thrusters is investigated for rate damping and maneuvers. The issues with implementing such an option are presented and the key issue of closed-loop stability is addressed. A new non-propulsive alternative to thruster control, Zero Propellant Maneuver (ZPM) attitude control method is introduced and its rate damping and maneuver performance evaluated. It is shown that ZPM can meet the tight attitude and rate error tolerances needed for long term attitude control. A combination of manual thruster rate damping to a safe harbor attitude followed by a ZPM to Stage long term attitude control orientation was selected by the Anomaly Resolution Team as the alternate attitude control method for such a contingency.
Pulse motor for spacecraft attitude control.
Multipulse thruster module design and development tests for spacecraft attitude control, noting micropulse solid propellant rocket motor
Restoring Redundancy to the MAP Propulsion System
The Microwave Anisotropy Probe (MAP) is a follow-on to the Differential Microwave Radiometer (DMR) instrument on the Cosmic Background Explorer (COBE). Due to the MAP project's limited mass, power, and financial resources, a traditional reliability concept including fully redundant components was not feasible. The MAP design employs selective hardware redundancy, along with backup software modes and algorithms, to improve the odds of mission success. In particular, MAP's propulsion system, which is used for orbit maneuvers and momentum management, uses eight thrusters positioned and oriented in such a way that its thruster-based attitude control modes can maintain three-axis attitude control in the event of the failure of any one thruster.
Torque equilibrium attitude control for Skylab reentry
The method of torque equilibrium attitude control used to control the reentry of Skylab to an altitude below 150 km without the use of thruster fuel once the attitude was established is discussed. The Skylab attitude and pointing control system, which included rate gyros, sun sensors, star tracker, the Apollo telescope mount digital computer, control moment gyros and cold-gas attitude thrusters, is presented. The 12 torque equilibrium attitudes found at which aerodynamic, gravity gradient and gyroscopic torques would balance are indicated, and the three of those at which the solar power supply would be adequate for attitude control are illustrated. The equilibrium seeking method employed is then examined, and the operation and performance of the torque equilibrium attitude control system during the three weeks prior to Skylab reentry are discussed. It is concluded that the torque equilibrium attitude control method developed for Skylab was successful in performing its assigned mission, and will be valuable for the design of future, low-altitude spacecraft or tethered vehicles.
Application of Micro-thrusters for Space Observatory Precision Attitude Control
This paper describes the results of a NASA investigation into the benefits of micro-thrusters compared to reaction wheels on future observatory-class missions with tight pointing stability requirements. Pointing repeatability and stability (i.e., jitter) requirements are key for space telescope missions of the future. For example, managing jitter is essential to being able to “image” planets orbiting distant stars. Jitter requirements for missions in this class are difficult to meet with current reaction wheel-based architectures. The reaction wheels are typically the largest pointing disturbance on the spacecraft. Disturbances from reaction wheels can be mitigated, typically by mechanically isolating the wheels, which imposes system complexity and cost. Thrusters capable of thrust forces in the micronewton (μN) range (referred to as micro-thrusters or micronewton thrusters) have been developed to support the Laser Interferometer Space Antenna (LISA) mission, which requires drag-free control to place a test mass in near-perfect free-fall Beyond the drag-free control application, micro-thrusters could be used as a substitute for reaction wheels or as a supplement to wheels for fine pointing control. Used in this fashion, micro-thrusters have potential for reducing the cost and technical risks of achieving demanding pointing stability performance on observatory-class missions.
Study of monopropellants for electrothermal thrusters: Design and fabrication task summary report
The feasibility of operating small thrust level electrothermal thrusters with monopropellants other than MIL-grade hydrazine was studied. Analytical study, design, and fabrication of demonstration thrusters was performed, and an evaluation test program was initiated to evaluate monopropellants with freezing points lower than MIL-grade hydrazine, and to determine their applicability to electrothermal thrusters for spacecraft attitude control. Five demonstration thrusters were fabricated to determine the feasibility of operation with monomethylhydrazine, Aerozine-50, 77 percent hydrazine-23 percent hydrazine azide, and a mixture of hydrazine monopropellants consisting of 35 percent hydrazine-50 percent monomethylhydrazine-15 percent ammonia. The present thruster is designed to produce a steady-state thrust level of 0.344 N at 1.724 x 1 million N/M sq feed pressure. Vacuum specific impulse goals were set at 1961 N-s/kg steady-state and 1716 N-s/kg pulsed-mode.
Coilgun Acceleration Model Containing Interactions Between Multiple Coils
Electromagnetic (EM) accelerators have the potential to fill a performance range not currently being met by conventional chemical and electric propulsion systems by providing a specific impulse of 600-1000 seconds and a thrust-to-power ratio greater than 200 mN/kW. A propulsion system based on EM acceleration of small projectiles has the traditional advantages of using a pulsed system, including precise control over a range of thrust and power levels as well as rapid response and repetition rates. Furthermore, EM accelerators have lower power requirements than conventional electric propulsion systems since no plasma creation is necessary. A coilgun is a specific type of EM device where a high-current pulse through a coil of wire interacts with a conductive projectile via an induced magnetic field to accelerate the projectile. There are no physical or electrical connections to the projectile, which leads to less system degradation and a longer life expectancy. Multi-staging a coilgun by adding multiple turns on a single coil or on the projectile increases the inductance, thus permitting acceleration of the projectile to higher velocities. Previously, a simplified problem of modeling an inductively-coupled, single-coil coilgun using a circuit-based analysis coupled to the one-dimensional momentum equation through Lenz's law was solved; however, the analysis was only conducted on uncoupled coils. The problem is significantly more complicated when multiple, independently-powered coils simultaneously operate and interact with each other and the projectile through induced magnetic fields. This paper presents a multi-coil model developed with the magnetostatic finite element solver QuickField. In the model, mutual inductance values between pairs of conductors were found by first computing the magnetic field energy for different cases where individual coils or multiple coils carry current, then integrating over the entire finite element domain for each case, and finally using the definition of inductive energy storage to solve for the self and mutual inductance. The electric circuit model is coupled to the projectile through Lenz's law, with the coils coupled through mutual inductance but able to be independently triggered at different times to optimize the acceleration profile. This initial model to predict the behavior of a projectile's acceleration through a coupled, multi-coil coilgun increases the potential of building a highly efficient coilgun thruster with key advantages over other EM thruster systems, thus making it a promising candidate for satellite main propulsion or attitude control thrusters.
Aerothermodynamics of the Mars Global Surveyor Spacecraft
The aerothermodynamics characteristics of the Mars Global Surveyor spacecraft are investigated and reported. These results have been used by the Mars Global Surveyor mission planners to design the aerobraking phase of the mission. Analytical and Direct Simulation Monte Carlo computer codes were used with a detailed, three dimensional model of the spacecraft to evaluate spacecraft aerobraking characteristics for flight in free molecular and transitional flow regimes. The spacecraft is found to be aerodynamically stable in aerobraking and planned contingency configurations. Aerodynamic forces, moments, and heating are found to be highly dependent on atmospheric density. Accommodation coefficient. is seen to strongly influence drag coefficient. Transitional flow effects are found to reduce overall solar panel heating. Attitude control thruster plumes are shown to interact with the freestream, diminishing the effectiveness of the attitude control system and even leading to thrust reversal. These plume-freestream interaction effects are found to be highly dependent on freestream density.