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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 343 records · Page 19

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NASA NTRS1981

Free radical propulsion concept

The concept of a free radical propulsion system, utilizing the recombination energy of dissociated low molecular weight gases to produce thrust, is analyzed. The system, operating at a theoretical impulse with hydrogen, as high as 2200 seconds at high thrust to power ratio, is hypothesized to bridge the gap between chemical and electrostatic propulsion capabilities. A comparative methodology is outlined by which characteristics of chemical and electric propulsion for orbit raising mission can be investigated. It is noted that free radicals proposed in rockets previously met with difficulty and complexity in terms of storage requirements; the present study proposes to eliminate the storage requirements by using electric energy to achieve a continuous-flow product of free radicals which are recombined to produce a high velocity propellant. Microwave energy used to dissociate a continuously flowing gas is transferred to the propellant via three-body-recombination for conversion to propellant kinetic energy. Microwave plasma discharge was found in excess of 90 percent over a broad range of pressure in preliminary experiments, and microwave heating compared to electrothermal heating showed much higher temperatures in gasdynamic equations.

Hawkins, C. E.↗

NASA NTRS1984

LOX/LH2 propulsion system for launch vehicle upper stage, test results

The test results of small LOX/LH2 engines for two propulsion systems, a pump fed system and a pressure fed system are reported. The pump fed system has the advantages of higher performances and higher mass fraction. The pressure fed system has the advantages of higher reliability and relative simplicity. Adoption of these cryogenic propulsion systems for upper stage of launch vehicle increases the payload capability with low cost. The 1,000 kg thrust class engine was selected for this cryogenic stage. A thrust chamber assembly for the pressure fed propulsion system was tested. It is indicated that it has good performance to meet system requirements.

Ikeda, T.↗

NASA NTRS2014

Multi-Objective Hybrid Optimal Control for Interplanetary Mission Planning

Preliminary design of low-thrust interplanetary missions is a highly complex process. The mission designer must choose discrete parameters such as the number of flybys, the bodies at which those flybys are performed, and in some cases the final destination. Because low-thrust trajectory design is tightly coupled with systems design, power and propulsion characteristics must be chosen as well. In addition, a time-history of control variables must be chosen which defines the trajectory. There are often may thousands, if not millions, of possible trajectories to be evaluated. The customer who commissions a trajectory design is not usually interested in a point solution, but rather the exploration of the trade space of trajectories between several different objective functions. This can be a very expensive process in terms of the number of human analyst hours required. An automated approach is therefore very desirable. This work presents such an approach by posing the mission design problem as a multi-objective hybrid optimal control problem. The method is demonstrated on hypothetical mission to the main asteroid belt and to Deimos.

Low Thrust↗

NASA NTRS2015

Multi-Objective Hybrid Optimal Control for Interplanetary Mission Planning

Preliminary design of low-thrust interplanetary missions is a highly complex process. The mission designer must choose discrete parameters such as the number of flybys, the bodies at which those flybys are performed, and in some cases the final destination. Because low-thrust trajectory design is tightly coupled with systems design, power and propulsion characteristics must be chosen as well. In addition, a time-history of control variables must be chosen which defines the trajectory. There are often many thousands, if not millions, of possible trajectories to be evaluated. The customer who commissions a trajectory design is not usually interested in a point solution, but rather the exploration of the trade space of trajectories between several different objective functions. This can be very expensive process in terms of the number of human analyst hours required. An automated approach is therefore very desirable. This work presents such an approach by posing the mission design problem as a multi-objective hybrid optimal control problem. The methods is demonstrated on hypothetical mission to the main asteroid belt and to Deimos.

Trajectory Design↗

NASA NTRS1968

Low variable thrust interplanetary trajectory data

Algorithm for generating low variable thrust interplanetary trajectory data for mission analysis of electrical propulsion systems

Mac Kay, J. S.↗

NASA NTRS2011

Low-Thrust Orbit Transfer Design for DAWN Operations at VESTA

Upon arrival at the asteroid Vesta, scheduled for July of 2011, the Dawn spacecraft will target a series of four distinct mapping orbits, each providing a unique opportunity to observe Vesta. The unknown, and potentially complex, Vesta gravity field presents challenges for designing low-thrust transfers between these mapping orbits while maintaining spacecraft safety from Vesta occultation of the Sun. This paper provides a description of the orbit transfers designed for Vesta operations along with a discussion of the constraints and methods used to design these transfers. The effect of alternate gravity fields on the viability of the designs and the design method is also considered

low thrust↗

NASA NTRS2014

Interplanetary Trajectory Design for the Asteroid Robotic Redirect Mission Alternate Approach Trade Study

This paper presents mission performance analysis methods and results for the Asteroid Robotic Redirect Mission (ARRM) option to capture a free standing boulder on the surface of a 100 m or larger NEA. It details the optimization and design of heliocentric low-thrust trajectories to asteroid targets for the ARRM solar electric propulsion spacecraft. Extensive searches were conducted to determine asteroid targets with large pick-up mass potential and potential observation opportunities. Interplanetary trajectory approximations were developed in method based tools for Itokawa, Bennu, 1999 JU3, and 2008 EV5 and were validated by end-to-end integrated trajectories.

Merrill, Raymond Gabriel↗

NASA NTRS2015

Low-Mass, Low-Power Hall Thruster System

NASA is developing an electric propulsion system capable of producing 20 mN thrust with input power up to 1,000 W and specific impulse ranging from 1,600 to 3,500 seconds. The key technical challenge is the target mass of 1 kg for the thruster and 2 kg for the power processing unit (PPU). In Phase I, Busek Company, Inc., developed an overall subsystem design for the thruster/cathode, PPU, and xenon feed system. This project demonstrated the feasibility of a low-mass power processing architecture that replaces four of the DC-DC converters of a typical PPU with a single multifunctional converter and a low-mass Hall thruster design employing permanent magnets. In Phase II, the team developed an engineering prototype model of its low-mass BHT-600 Hall thruster system, with the primary focus on the low-mass PPU and thruster. The goal was to develop an electric propulsion thruster with the appropriate specific impulse and propellant throughput to enable radioisotope electric propulsion (REP). This is important because REP offers the benefits of nuclear electric propulsion without the need for an excessively large spacecraft and power system.

Pote, Bruce↗

NASA NTRS2017

Validation of a Low-Thrust Mission Design Tool Using Operational Navigation Software

Design of flight trajectories for missions employing solar electric propulsion requires a suitably high-fidelity design tool. In this work, the Evolutionary Mission Trajectory Generator (EMTG) is presented as a medium-high fidelity design tool that is suitable for mission proposals. EMTG is validated against the high-heritage deep-space navigation tool MIRAGE, demonstrating both the accuracy of EMTG's model and an operational mission design and navigation procedure using both tools. The validation is performed using a benchmark mission to the Jupiter Trojans.

Englander, Jacob A.↗

NASA NTRS2017

Validation of a Low-Thrust Mission Design Tool Using Operational Navigation Software

Englander, Jacob A.↗

NASA NTRSDate not supplied

Preliminary Statistical Maneuver Analysis for A Low-Thrust NRHO to DRO Transfer

To understand practical implications of implementing transfers that leverage multibody dynamics, a statistical analysis is performed on an Earth-Moon Near Rectilinear Halo Orbit (NRHO) to distant retrograde orbit (DRO) transfer using solar electric propulsion. The statistical analysis includes perturbations due to maneu- ver execution and orbit determination error to understand operational character- istics of the transfer such as statistical ∆V and statistical state targeting errors at the final DRO. We explore the need for low-thrust trajectory correction maneuver (TCM) opportunities as well as the delivery accuracy to the final target. Finally, we explore options for long-term-stable target DROs.

low thrust↗

NASA NTRS2004

The NASA GSFC MEMS Colloidal Thruster

A number of upcoming missions require different thrust levels on the same spacecraft. A highly scaleable and efficient propulsion system would allow substantial mass savings. One type of thruster that can throttle from high to low thrust while maintaining a high specific impulse is a Micro-Electro-Mechanical System (MEMS) colloidal thruster. The NASA GSFC MEMS colloidal thruster has solved the problem of electrical breakdown to permit the integration of the electrode on top of the emitter by a novel MEMS fabrication technique. Devices have been successfully fabricated and the insulation properties have been tested to show they can support the required electric field. A computational finite element model was created and used to verify the voltage required to successfully operate the thruster. An experimental setup has been prepared to test the devices with both optical and Time-Of-Flight diagnostics.

Cardiff, Eric H.↗

NASA NTRSDate not supplied

Mission Design Considerations for a Low-Thrust Spacecraft

Developing an executable low-thrust trajectory for use in a spaceflight mission requires the design and optimization of a deterministic trajectory as well as the validation that the selected architecture is robust to some set of uncertainties, execution errors, and potential contingencies. Uncertainty in the ability of the spacecraft and its launch vehicle to execute a trajectory as well as potential deviations such as in-flight anomalies combine with design-to constraints and requirements to complicate the optimization problem. The approach by which robust mission design was accomplished for the initial capability of NASA’s Gateway is presented as well as associated results.

mission design↗

NASA NTRSDate not supplied

Recovery from Missed Thrust During Low Thrust Insertion of NASA's Gateway into a Near Rectilinear Halo Orbit

Various strategies are assessed by which the first two elements of Gateway could recover from missed thrust events during insertion into a Near Rectilinear Halo Orbit (NRHO). Recovery strategies are assessed in terms of their ability to maintain the desired NRHO insertion epoch ad well as their cost in terms of additional time and ∆v required to reach the NRHO. Strategies utilize a 50 kW Solar Electric Propulsion (SEP) system onboard the Power and Propulsion Element (PPE). Recovery solutions are found that, in the event of a missed thrust arc during the Insertion Phase of the Gateway Lunar Transit, deliver the space-craft into the NRHO within three additional revolutions (roughly 21 days) and 80 m/s ∆v.

low thrust↗

NASA NTRSDate not supplied

Recovery from Missed Thrust During Low Thrust Insertion of NASA's Gateway into a Near Rectilinear Halo Orbit

low thrust↗

NASA NTRS1991

Sunmaster: An SEP cargo vehicle for Mars missions

Options are examined for an unmanned solar powered electric propulsion cargo vehicle for Mars missions. The 6 prime areas of study include: trajectory, propulsion system, power system, supporting structure, control system, and launch consideration. Optimization of the low thrust trajectory resulted in a total round trip mission time just under 4 years. The argon propelled electrostatic ion thruster system consists of seventeen 5 N engines and uses a specific impulse of 10,300 secs. At Earth, the system uses 13 engines to produce 60 N of thrust; at Mars, five engines are used, producing 25 N thrust. The thrust of the craft is varied between 60 N at Earth and 24 N at Mars due to reduced solar power available. Solar power is collected by a Fresnel lens concentrator system using a multistacked cell. This system provides 3.5 MW to the propulsion system after losses. Control and positioning to the craft are provided by a system of three double gimballed control moment gyros. Four shuttle 'C' launches will be used to transport the unassembled vehicle in modular units to low Earth orbit where it will be assembled using the Mobile Transporter of the Space Station Freedom.

Chiles, Aleasa↗

NASA NTRS1998

Demonstrating the Performance Benefits of the Strutjet RBCC for Space Launch Architectures

The Rocket Based Combined Cycle (RBCC) engine synergistically combines the best elements of airbreathing and rocket propulsion to benefit a wide range of future reusable launch vehicles. Aerojet's Strutjet RBCC offers high I(sub sp) during mid-phase acceleration, and high thrust for boost and final ascent phases. The result is a relatively low gross weight vehicle with lower thrust requirements than comparable all-rocket solutions. Relative to combination propulsion systems, the integrated propulsive elements of the Strutjet reduce engine weight and complexity. In the recent NASA Highly Reusable Space Transportation (HRST) study, vehicles with RBCC-class engines were projected to have lower per flight operations costs than comparable rockets. The more benign operating environment of the RBCC engine primary rockets (2000 psi chamber pressure, lower internal temperatures, gas generator cycle) lead to longer service life, fewer maintenance actions, and increased reliability. Operations costs are typically the most significant contributor to life cycle costs. RBCC-powered vehicles lend themselves naturally to horizontal takeoff, which typically requires less thrust for a given gross vehicle weight when compared to an all-rocket-powered vehicle. The lower thrust requirements reduce engine size and/or quantity, reducing the n of ground personnel and facilities required to maintain the vehicle. The reduced sensitivity to mass fraction also allows the vehicle to be designed with higher margins for improved life and reliability. Because of its multi-mode ascent operation and ability to perform long cruise segments, the strutjet RBCC offers RLV designers a chance to create vehicles with an unprecedented level of mission flexibility. Features such as supersonic cruise, flyback, off-set launch, self-ferry, horizontal takeoff and landing, multi-base operation, and flexible aborts open up a broad range of potential revenue producing missions. The Strutjet can lead to low-cost vehicle designs in traditional areas such as LEO cargo delivery and ISS resupply, and also lead to economically attractive vehicles for small payload missions difficult to achieve with all-rocket solutions. The performance capability of the Strutjet RBCC has been demonstrated in all operating mod underfunding from NASA MSFC and Aerojet as a result of over 1000testsand over 13 years of continuous development-a significant resource investment in the Strutjet engine design.

Johnson, D.↗

NASA NTRS1974

HILTOP supplement: Heliocentric interplanetary low thrust trajectory optimization program, supplement 1

Modifications and improvements are described that were made to the HILTOP electric propulsion trajectory optimization computer program during calendar years 1973 and 1974. New program features include the simulation of power degradation, housekeeping power, launch asymptote declination optimization, and powered and unpowered ballistic multiple swingby missions with an optional deep space burn.

Mann, F. I.↗