Thruster Requirements Research Articles

DESIGN AND PERFORMANCE OF A 110 N ATTITUDE CONTROL THRUSTER FOR LUNAR MISSIONS

The design and development of an attitude control thruster, designated A110, developed for broad attitude control applications, including lunar landers, is described. The A110 thruster produces 110 N (25 lbf) of nominal thrust and in-excess of 305.5 s of specific impulse (I<sub>sp</sub>) while operating on a fuel blend of 20% monomethylhydrazine-80% hydrazine (M20) with mixed oxides of nitrogen containing 3% nitric oxide (MON3) as the oxidizer. The injector and combustion chamber are additively manufactured with cobalt-chrome and niobium C103, respectively, to enable the rapid development required of the industry. The thruster is capable of response times of under 10 ms and currently produces a minimum impulse bit (MIB) of 0.436 N.s and has demonstrated operation in pulse-mode with a variety of duty cycles during development testing. These capabilities allow for the generation of precise pulses for accurate control of a lander. The testing of this thruster is used to evaluate steady-state and pulse-mode specific impulse, impulse repeatability, and combustion stability. The measured performance (i.e., specific impulse and characteristic velocity (C*), are compared with theoretical values. The capabilities demonstrated by the A110 thruster are applicable to a wide range of lunar lander missions and can be qualified for minor changes in thruster requirements for additional missions without the need for major design changes. This article is published with the permission of the authors granted to 3AF - Association Aéronautique et Astronautique de France (www.3AF.fr) organizer of the Space Propulsion International Conference.

Astrodynamical middle-frequency interferometric gravitational wave observatory AMIGO: Mission concept and orbit design

AMIGO is a first-generation Astrodynamical Middle-frequency Interferometric Gravitational Wave (GW) Observatory. The scientific goals of AMIGO are to bridge the spectra gap between first-generation high-frequency and low-frequency GW sensitivities: to detect intermediate mass BH coalescence; to detect inspiral phase and predict time of binary black hole coalescences together with binary neutron star & black hole-neutron star coalescences for ground interferometers; to detect compact binary inspirals for studying stellar evolution and galactic population. The mission concept is to use time delay interferometry (TDI) for a nearly triangular formation of three drag-free spacecraft with nominal arm length 10,000 km, emitting laser power 2–10 W and telescope diameter 300–500 mm. The design GW sensitivity in the middle frequency band is [Formula: see text] Hz[Formula: see text]. Both geocentric and heliocentric orbit formations are considered. All options have LISA-like formations, that is, the triangular formation is [Formula: see text] inclined to the orbit plane. For all solar orbit options of AMIGO, the first-generation TDI satisfies the laser frequency-noise suppression requirement. We also investigate for each option of orbits under study, whether constant equal-arm implementation is feasible. For the solar-orbit options, the acceleration to maintain the formation can be designed to be less than 15 nm/s2with the thruster requirement in the 15 [Formula: see text]N range. AMIGO would be a good place to test the feasibility of the constant equal-arm option. Fuel requirement, thruster noise requirement and test mass acceleration actuation requirement are briefly considered. From the orbit study, the solar orbit option is the mission orbit preference. We study the deployment for this orbit option. After a last-stage launch from 300 km Low Earth Orbit (LEO), each S/C’s maneuver to an appropriate 2-degree-behind-the-Earth AMIGO formation in 95 days requires only a [Formula: see text]v of about 80 m/s.

Orbit design and thruster requirement for various constant arm space mission concepts for gravitational-wave observation

In previous papers, we have addressed the issues of orbit design and thruster requirement for the constant arm versions of Astrodynamical Middle-frequency Interferometric Gravitational-wave Observatory (AMIGO) mission concept and for the constant arm gravitational wave (GW) mission concept of Atom Interferometric Gravitational-wave Space Observatory (AIGSO). In this paper, we apply similar methods to the orbit design and thruster requirement for the constant arm GW missions B-DECIGO and DECIGO, and estimate the yearly propellant requirements at the specific impulse [Formula: see text][Formula: see text]s and [Formula: see text][Formula: see text]s. For the geocentric orbit options of B-DECIGO which we have explored, the fuel mass requirement is a concern. For the heliocentric orbit options of B-DECIGO and DECIGO, the fuel requirement to keep the arm equal and constant should be easily satisfied. Furthermore, we explore the thruster and propellant requirements for constant arm versions of LISA and TAIJI missions and find the fuel mass requirement is not a show stopper either. The proof mass actuation noise is a concern. To have enough dynamical range, an alternate proof mass is required. Detailed laboratory study is warranted.

Orbit design for space atom-interferometer AIGSO

Atom-interferometric gravitational-wave (GW) space observatory (AIGSO) is a mission concept mainly aimed at the middle-frequency (0.1-10 Hz) GW detection. AIGSO proposes to have three spacecrafts in linear formation with extension of 10[Formula: see text]km. The three spacecrafts need to maintain [Formula: see text] constant arm-length formation. In this study, we address the issue of orbit design and thruster requirement. The acceleration to maintain the formation can be designed to be less than 30[Formula: see text]pm/s2and the thruster requirement is in the 30[Formula: see text]nN range. Application to other arm-length-maintaining missions is also discussed.

Engineering Risk Assessment of a dynamic space propulsion system benchmark problem

The Engineering Risk Assessment (ERA) team at NASA Ames Research Center develops dynamic models with linked physics-of-failure analyses to produce quantitative risk assessments of space exploration missions. This paper applies the ERA approach to the 2014 Probabilistic Safety Assessment and Management conference Space Propulsion System Benchmark Problem, which investigates dynamic system risk for a deep space ion propulsion system over three missions with time-varying thruster requirements and operations schedules. The dynamic missions are simulated using commercial software to generate integrated loss-of-mission (LOM) probability results via Monte Carlo sampling. The simulation model successfully captured all dynamics aspects of the benchmark missions, and convergence studies are presented to illustrate the sensitivity of integrated LOM results to the number of Monte Carlo trials. In addition, to evaluate the relative importance of dynamic modeling, the Ames Reliability Tool (ART) was used to build a series of quasi-dynamic, deterministic models that incorporated varying levels of the problem׳s dynamics. The ART model did a reasonable job of matching the simulation results for the simpler mission case, while auxiliary dynamic models were required to adequately capture risk-driver rankings for the more dynamic cases. This study highlights how state-of-the-art techniques can adapt to a range of dynamic problems.

Micronewton Thruster Requirements for Earth-Orbiting Imaging Formations

Previous studies have explored options for separated spacecraft interferometers for planet detection as well as optical imaging of astrophysical objects. As others have noted, one of the benefits of Earth-orbiting formations is the potential to reduce the propellant required to complete image (u-v) plane filling maneuvers by taking advantage of gravity-gradient forces. In this work, the equations of motion for each of the three spacecraft in a linear formation are numerically integrated using an acceleration profile previously proposed by DeCou to produce a gradual filling of the image plane. The analysis also includes the addition of the J 2 perturbing acceleration and provides the detailed spacecraft position, velocity, and thrust histories for a maneuver as well as the maximum baseline rate of change and ΔV. Results suggest that micronewton thrusters under development for other missions requiring precision spacecraft control might be enabling for the Earth-orbiting imaging formation considered here. For one of the reference cases in geostationary orbit with a maximum baseline of 1 km and an allocated maneuver time of five days, the maximum thrust and throttling rate are approximately 87 pN and 0.24 μN/min, respectively, with a maximum baseline rate of change of 2.4 cm/s and ΔV of 6 cm/s for a single maneuver (one observation). The same reference case with thruster cancellation of the J 2 acceleration results in a maximum thrust of 2.1 mN and ΔV of 3.6 The The maximum thrust, baseline rate of change, and ΔV are presented for maneuver times of one to five days, maximum baselines of 250 m and 1 km, and a formation altitude of 35,786 and 20,000 km. Three propulsion technology options-pulsed plasma, field emission electric propulsion, and colloid-are discussed as possible candidates for this role. Two solution strategies are proposed to the problem raised by the J 2 perturbation.

Thruster requirements and concerns for bipropellant blowdown systems

Bipropellant blowdown applications can involve a wide range of operating conditions and requirements for thrusters in the system. These requirements are determined by studies and analyses plus computer simulations of propulsion system operation. These studies, in conjunction with associated thruster testing, have resulted in the identification of significant areas of concern that should be addressed, including performance effects, chamber life, ignition spikes, and combustion stability. There are also potential duty cycle sensitivities, and successful operation in the steady-state mode does not guarantee acceptable thermal control during pulsing modes. Testing should be conducted using realistic inlet conditions and duty cycles, and caution should be exercised in attempting to extrapolate into regions that have not been covered by actual testing.

Power conditioning for space propulsion systems using contact ion thrusters.

The paper is concerned with the problems associated with matching the output of a large nuclear- electric power source to the requirements of contact ion thrusters. The factors that influence the design requirements and the technical approach to power conditioning, switching, and control (PCSC) subsystems are discussed. Thruster requirements, power source characteristics, and environmental constraints are analyzed from the standpoint of their relationships to power conditioning. In addition, the results of the development of a breadboard PCSC unit at a 30-kw power level are described. This work gives an insight into the problems and the possible solutions that will result in the realization of power conditioning equipment whose specific weight, efficiency, and reliability are compatible with requirements of large scale nuclear- electric rockets of the future. In particular, one of the most promising experimental results is a 30-kw breadboard whose components exhibit a total weight of 3.2 lb/ kw. The results of a design study for a 300-kw system show that it is feasible to build a PCSC system weighing 9.1 Ib/kw input power including radiators and shadow shielding with a power efficiency of 93%.