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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 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

On the calculation of low-thrust fail-safe trajectories

A guidance algorithm is developed for a low-thrust spacecraft such that target intercept is possible in spite of premature thrust termination along the trajectory. Such a trajectory is called a 'fail-safe' trajectory and the spacecraft thrust is utilized to minimize the relative target-spacecraft approach speed. The fail-safe guidance algorithm is solved using the concept of a critical thrust plane and a non-critical thrust direction. Several examples of fail-safe guidance are presented for a solar-electric propulsion flyby mission to the comet Encke.

Sauer, C. G., Jr.↗
NASA NTRS1980

LEO-to-GEO low thrust chemical propulsion

One approach being considered for transporting large space structures from low Earth orbit (LEO) to geosynchronous equatorial orbit (GEO) is the use of low thrust chemical propulsion systems. A variety of chemical rocket engine cycles evaluated for this application for oxygen/hydrogen and oxygen/hydrocarbon propellants (oxygen/methane and oxygen/RF-1) are discussed. These cycles include conventional propellant turbine drives, turboalternator/electric motor pump drive, and fuel cell/electric motor pump drive as well as pressure fed engines. Thrust chamber cooling analysis results are presented for regenerative/radiation and film/radiation cooling.

Shoji, J. M.↗
NASA NTRS2017

A Time-Regularized Multiple Gravity-Assist Low-Thrust Bounded-Impulse Model for Trajectory Optimization

The multiple gravity assist low-thrust (MGALT) trajectory model combines the medium-fidelity Sims-Flanagan bounded-impulse transcription with a patched-conics flyby model and is an important tool for preliminary trajectory design. While this model features fast state propagation via Kepler's equation and provides a pleasingly accurate estimation of the total mass budget for the eventual flight-suitable integrated trajectory it does suffer from one major drawback, namely its temporal spacing of the control nodes. We introduce a variant of the MGALT transcription that utilizes the generalized anomaly from the universal formulation of Kepler's equation as a decision variable in addition to the trajectory phase propagation time. This results in two improvements over the traditional model. The first is that the maneuver locations are equally spaced in generalized anomaly about the orbit rather than time. The second is that the Kepler propagator now has the generalized anomaly as its independent variable instead of time and thus becomes an iteration-free propagation method. The new algorithm is outlined, including the impact that this has on the computation of Jacobian entries for numerical optimization, and a motivating application problem is presented that illustrates the improvements that this model has over the traditional MGALT transcription.

Optimization↗
NASA NTRS2017

A Time-Regularized, Multiple Gravity-Assist Low-Thrust, Bounded-Impulse Model for Trajectory Optimization

The multiple gravity assist low-thrust (MGALT) trajectory model combines the medium-fidelity Sims-Flanagan bounded-impulse transcription with a patched-conics flyby model and is an important tool for preliminary trajectory design. While this model features fast state propagation via Keplers equation and provides a pleasingly accurate estimation of the total mass budget for the eventual flight suitable integrated trajectory it does suffer from one major drawback, namely its temporal spacing of the control nodes. We introduce a variant of the MGALT transcription that utilizes the generalized anomaly from the universal formulation of Keplers equation as a decision variable in addition to the trajectory phase propagation time. This results in two improvements over the traditional model. The first is that the maneuver locations are equally spaced in generalized anomaly about the orbit rather than time. The second is that the Kepler propagator now has the generalized anomaly as its independent variable instead of time and thus becomes an iteration-free propagation method. The new algorithm is outlined, including the impact that this has on the computation of Jacobian entries for numerical optimization, and a motivating application problem is presented that illustrates the improvements that this model has over the traditional MGALT transcription.

Optimization↗
NASA NTRS2019

Efficient computation of optimal low thrust gravity perturbed orbit transfers

We have developed a new method for solving low-thrust fuel-optimal orbit transfer problems in the vicinity of a large body (planet or asteroid), considering a high-fidelity spherical harmonic gravity model. The algorithm is formulated via the indirect optimization method, leading to a two-point boundary value problem (TPBVP). We make use of a hyperbolic tangent smoothing law for performing continuation on the thrust magnitude to reduce the sharpness of the control switches in early iterations and thus promote convergence. The TPBVP is solved using the method of particular solutions (MPS) shooting method and Picard-Chebyshev numerical integration. Application of Picard-Chebyshev integration affords an avenue for increased efficiency that is not available with step-by-step integrators. We demonstrate that computing the particular solutions with only a low-fidelity force model greatly increases the efficiency of the algorithm while ultimately achieving near machine precision accuracy. A salient feature of the MPS is that it is parallelizable, and thus further speedups are available. It is also shown that, for near Earth orbits and over a small number of en-route revolutions around the Earth, only the zonal perturbation terms are required in the costate equations to obtain a solution that is accurate to machine precision and optimal to engineering precision. The proposed framework can be used for trajectory design around small asteroids and also for orbit debris rendezvous and removal tasks.

Junkins, John L.↗
NASA NTRS1991

Low thrust viscous nozzle flow fields prediction

A Navier-Stokes code was developed for low thrust viscous nozzle flow field prediction. An implicit finite volume in an arbitrary curvilinear coordinate system lower-upper (LU) scheme is used to solve the governing Navier-Stokes equations and species transportation equations. Sample calculations of carbon dioxide nozzle flow are presented to verify the validity and efficiency of this code. The computer results are in reasonable agreement with the experimental data.

Liaw, G. S.↗
NASA NTRS1975

Limited variance control in statistical low thrust guidance analysis

Difficulties arise in guiding a solar electric propulsion spacecraft due to nongravitational accelerations caused by random fluctuations in the magnitude and direction of the thrust vector. These difficulties may be handled by using a low thrust guidance law based on the linear-quadratic-Gaussian problem of stochastic control theory with a minimum terminal miss performance criterion. Explicit constraints are imposed on the variances of the control parameters, and an algorithm based on the Hilbert space extension of a parameter optimization method is presented for calculation of gains in the guidance law. The terminal navigation of a 1980 flyby mission to the comet Encke is used as an example.

Jacobson, R. A.↗
NASA NTRS1981

Low-thrust chemical rocket engine study

An analytical study evaluating thrust chamber cooling engine cycles and preliminary engine design for low thrust chemical rocket engines for orbit transfer vehicles is described. Oxygen/hydrogen, oxygen/methane, and oxygen/RP-1 engines with thrust levels from 444.8 N to 13345 N, and chamber pressures from 13.8 N/sq cm to 689.5 N/sq cm were evaluated. The physical and thermodynamic properties of the propellant theoretical performance data, and transport properties are documented. The thrust chamber cooling limits for regenerative/radiation and film/radiation cooling are defined and parametric heat transfer data presented. A conceptual evaluation of a number of engine cycles was performed and a 2224.1 N oxygen/hydrogen engine cycle configuration and a 2224.1 N oxygen/methane configuration chosen for preliminary engine design. Updated parametric engine data, engine design drawings, and an assessment of technology required are presented.

Shoji, J. M.↗
NASA NTRS1980

Low-thrust chemical rocket engine study

Parametric data and preliminary designs on liquid rocket engines for low thrust cargo orbit-transfer-vehicles are described and those items where technology is required to enhance the designs are identified. The results of film cooling studies to establish the upper chamber pressure limit are given. The study showed that regen cooling with RP-1 was not feasible over the entire thrust and chamber pressure ranges. The thermal data showed that the RP-1 bulk temperature exceeded the study coking temperature limit of 1010 R. Based upon the results presented, O2/H2 and O2/CH4 regen engine systems and O2/H2 film cooled engines were selected for further study in the system analysis. Six engine design concepts are examined.

Mellish, J. A.↗
NASA NTRS1984

Evaluation of propellent tank insulation concepts for low-thrust chemical propulsion systems: Executive summary

Cryogenic propellant tank insulations or liquid oxygen/liquid hydrogen low-thrust 2224N (500 lbf) propulsion systems (LTPS) were assessed. The insulation studied consisted of combinations of N2-purged foam and multilayer insulation (MLI) as well as He-purged MLI-only. Heat leak and payload performance predictions were made for three shuttle-launched LTPS designed for shuttle bay packaged payload densities of 56 kg cu/m (3.5 lbm/cu ft), 40 kg/cu m (2.5 lbm/cu ft) and 24 kg/cu m (1.5 lbm/cu ft). Foam/MLI insulations were found to increase LTPS payload delivery capability when compared with He-purged MLI-only. An additional benefit of foam/MLI was reduced operational complexity because orbiter cargo bay N2 purge gas could be used for MLI purging. Maximum payload mass benefit occurred when an enhanced convection, rather than natural convection, heat transfer was specified for the insulation purge enclosure. The enhanced convection environment allowed minimum insulation thickness to be used for the foam/MLI interface temperature selected to correspond to the moisture dew point in the N2 purge gas. Experimental verification of foam/MLI benefits was recommended. A conservative program cost estimate for testing a MLI-foam insulated tank was 2.1 million dollars. This cost could be reduced significantly without increasing program risk.

Kramer, T.↗
NASA NTRS1985

Small centrifugal pumps for low thrust rockets

This paper presents the results of a combined analytical and experimental investigation of low specific speed pumps for potential use as components of propellant feed systems for low thrust rocket engines. Shrouded impellers and open face impellers were tested in volute type and vaned diffuser type pumps. Full- and partial-emission diffusers and full- and partial-admission impellers were tested. Axial and radial loads, head and efficiency versus flow, and cavitation tests were conducted. Predicted performance of two pumps are compared when pumping water and liquid hydrogen. Detailed pressure loss and parasitic power values are presented for two pump configurations. Partial-emission diffusers were found to permit use of larger impeller and diffuser passages with a minimal performance penalty. Normal manufacturing tolerances were found to result in substantial power requirement variation with only a small pressure rise change. Impeller wear ring leakage was found to reduce pump pressure rise to an increasing degree as the pump flowrate was decreased.

Gulbrandsen, N. C.↗
NASA NTRS1992

Low thrust chemical rocket technology

An on-going technology program to improve the performance of low thrust chemical rockets for spacecraft on-board propulsion applications is reviewed. Improved performance and lifetime is sought by the development of new predictive tools to understand the combustion and flow physics, introduction of high temperature materials and improved component designs to optimize performance, and use of higher performance propellants. Improved predictive technology is sought through the comparison of both local and global predictions with experimental data. Predictions are based on both the RPLUS Navier-Stokes code with finite rate kinetics and the JANNAF methodology. Data were obtained with laser-based diagnostics along with global performance measurements. Results indicate that the modeling of the injector and the combustion process needs improvement in these codes and flow visualization with a technique such as 2-D laser induced fluorescence (LIF) would aid in resolving issues of flow symmetry and shear layer combustion processes. High temperature material fabrication processes are under development and small rockets are being designed, fabricated, and tested using these new materials. Rhenium coated with iridium for oxidation protection was produced by the Chemical Vapor Deposition (CVD) process and enabled an 800 K increase in rocket operating temperature. Performance gains with this material in rockets using Earth storable propellants (nitrogen tetroxide and monomethylhydrazine or hydrazine) were obtained through component redesign to eliminate fuel film cooling and its associated combustion inefficiency while managing head end thermal soakback. Material interdiffusion and oxidation characteristics indicated that the requisite lifetimes of tens of hours were available for thruster applications. Rockets were designed, fabricated, and tested with thrusts of 22, 62, 440 and 550 N. Performance improvements of 10 to 20 seconds specific impulse were demonstrated. Higher performance propellants were evaluated: Space storable propellants, including liquid oxygen (LOX) as the oxidizer with nitrogen hydrides or hydrocarbon as fuels. Specifically, a LOX/hydrazine engine was designed, fabricated, and shown to have a 95 pct theoretical c-star which translates into a projected vacuum specific impulse of 345 seconds at an area ratio of 204:1. Further performance improvment can be obtained by the use of LOX/hydrogen propellants, especially for manned spacecraft applications, and specific designs must be developed and advanced through flight qualification.

Schneider, Steven J.↗
NASA NTRS2004

Low-thrust Orbit Transfers Using Candidate Lyapunov Functions with a Mechanism for Coasting

We consider low-thrust orbit transfers around a central body, where specified changes are sought in orbit elements except true anomaly. The desired changes in the remaining five elements can be arbitrarily large. Candidate Lyapunov functions are created based on analytic expressions for maximum rates of change of the orbit elements and the desired changes in the elements. These functions may be thought of as proximity quotients because they provide R measure of the proximity to the target orbit. The direction of thrust needed for steepest descent to the target orbit is also available analytically. The thrust is shutoff if the effectivity of the thrust at the current location on the osculating orbit is below some threshhold value. Thus, the equations of motion can be numerically integrated to obtain quickly and simply a transfer to the target orbit. A series of transfers can be easily computed to assess the trade-off between propellant mass and flight time. Preliminary comparisons to optimal solutions show that the method, while sub-optimal, performs well.

Candidate Lyapunoc functions↗