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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 271 records · Page 15

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NASA NTRS1999

1-D Numerical Analysis of ABCC Engine Performance

ABCC engine combines air breathing and rocket engine into a single engine to increase the specific impulse over an entire flight trajectory. Except for the heat source, the basic operation of the ABCC is similar to the basic operation of the RBCC engine. The ABCC is intended to have a higher specific impulse than the RBCC for single stage Earth to orbit vehicle. Computational fluid dynamics (CFD) is a useful tool for the analysis of complex transport processes in various components in ABCC propulsion system. The objective of the present research was to develop a transient 1-D numerical model using conservation of mass, linear momentum, and energy equations that could be used to predict flow behavior throughout a generic ABCC engine following a flight path. At specific points during the development of the 1-D numerical model a myriad of tests were performed to prove the program produced consistent, realistic numbers that follow compressible flow theory for various inlet conditions.

Holden, Richard↗

NASA NTRS1973

Apollo 16 LM-11 descent propulsion system final flight evaluation

The performance of the LM-11 Descent Propulsion System during the Apollo 16 Mission was evaluated and found to be satisfactory. The average engine effective specific impulse was 0.1 second higher than predicted, but well within the predicted 1 sigma uncertainty of 0.2 seconds. The engine performance corrected to standard inlet conditions for the FTP portion of the burn at 50 seconds after ignition was as follows: thrust, 9839 lbf; specific impulse, 306.9 sec; and propellant mixture ratio, 1.592. These values are +0.34, +0.03 and +0.0 percent different, respectively, from the values reported from engine acceptance tests and were within specification limits. Several flight measurement discrepancies that existed during the flight are discussed.

Avvenire, A. T.↗

NASA NTRS1973

An 8-cm electron bombardment thruster for auxiliary propulsion

Thrustor size, beam current level, and specific impulse tradeoffs are considered for mercury electron bombardment ion thrustors to be used for north-south station keeping of geosynchronous spacecraft. An 8-cm-diam thrustor operating at 2750 sec specific impulse at thrust levels of 4.4 mN (1 mlb) to 8.9 mN (2 mlb) with a design life of 20,000 hours and 10,000 cycles is being developed. The thrustor will have a dished two-grid system capable of thrust vectoring of plus or minus 10 deg in two orthogonal directions. A preliminary thrustor has been fabricated and tested; thrustor performance characteristics have been determined at 4.45, 6.68, and 8.90 millinewtons.

Hudson, W. R.↗

NASA NTRS1987

Transport vehicle for manned Mars missions powered by inertial confinement fusion

Inertial confinement fusion (ICF) is an ideal engine power source for manned spacecraft to Mars because of its inherently high power-to-mass ratios and high specific impulses. In this paper a concept is produced for a vehicle powered by ICF and utilizing a magnetic thrust chamber to avoid plasma thermalization with wall structures and the resultant degradation of specific impulse, that are unavoidable with the use of mechanical thrust chambers. This vehicle is capable of 100-day manned Mars missions with a 100-metric-ton payload and a total vehicle launch mass near 6000 metric tons, based on advanced technology assumed to be available by A.D. 2020.

Orth, Charles D.↗

NASA NTRS1973

An 8-cm electron bombardment thruster for auxiliary propulsion

Thruster size, beam current level, and specific impulse trade-offs are considered for mercury electron bombardment ion thrusters to be used for north-south station keeping of geosynchronous spacecraft. An 8-cm diameter thruster operating at 2750 seconds specific impulse at thrust levels of 4.4 mN (1 m1b) to 8.9 mN (2 m6b) with a design life of 20,000 hours and 10,000 cycles is being developed. The thruster will have a dished two-grid system capable of thrust vectoring of + or - 10 degrees in two orthogonal directions. A preliminary thruster has been fabricated and tested; thruster performance characteristics have been determined at 4.45, 6.68, and 8.90 millinewtons.

Hudson, W. R.↗

NASA NTRS1982

Extended-performance 8-cm ion thruster operation

The performance of an 8-cm ion thruster has been extended to 32-mN thrust, 4062-s specific impulse, and 0.033-mN/W thrust-to-power ratio. The design of this thruster is similar to the IAPS (Ion Auxiliary Propulsion System), which has a performance of 5-mN thrust, 3500-s specific impulse, and 0.036-mN/W thrust-to-power ratio. The above level of extended performance has been achieved by increasing the discharge power, ion-beam accelerating voltage, and propellant flowrate, and by modifying the propellant-vapor injection, discharge-chamber magnetic-field strength, high-voltage insulators, and electron baffle. This paper describes results obtained in characterizing the extended-performance operation of both the unmodified thruster (i.e., in an IAPS-like configuration) and the modified thruster. Lifetime-estimating measurements for the modified thruster at extended-performance conditions are also presented.

Wessel, F. J.↗

NASA NTRS1983

Development of advanced inert-gas ion thrusters

Inert gas ion thruster technology offers the greatest potential for providing high specific impulse, low thrust, electric propulsion on large, Earth orbital spacecraft. The development of a thruster module that can be operated on xenon or argon propellant to produce 0.2 N of thrust at a specific impulse of 3000 sec with xenon propellant and at 6000 sec with argon propellant is described. The 30 cm diameter, laboratory model thruster is considered to be scalable to produce 0.5 N thrust. A high efficiency ring cusp discharge chamber was used to achieve an overall thruster efficiency of 77% with xenon propellant and 66% with argon propellant. Measurements were performed to identify ion production and loss processes and to define critical design criteria (at least on a preliminary basis).

Poeschel, R. L.↗

NASA NTRSDate not supplied

One Million Newton-Second Duration Test of a 600 Watt Hall Effect Thruster Fueled By Xenon

A Hall Effect Thruster fueled by xenon was duration tested for7,198 hours at NASA’s Glenn Research Center.The thruster generated 1 million Newton-seconds of total impulse before the test was terminated voluntarily. Thrust, flow rate, and many other parameters were tracked continuously. Over 98.5% of the test was conducted at a single operating point, where the discharge power was 600 W, mean thrust was 39 mN and mean specific impulse was 1500 s. The shape of the discharge channel was measured via laser profilometery. The shape changed considerably over the course of the test, due to ion sputtering. However, the standard deviation of the thrust and specific impulse measurements at the nominal operating point was less than 1%. Evolution of the plasma plume was measured with a Faraday probe mounted upon a rotating arm. Plume divergence decreased over the course of the test

Electric Propulsion↗

NASA NTRS1978

Extended-performance thruster technology evaluation

Two 30-cm ion thruster technology areas are investigated in support of the extended-performance thruster operation required for the Halley's comet rendezvous mission. These areas include an evaluation of the thruster performance and lifetime characteristics at increased specific impulse and power levels, and the design and evaluation of a high-voltage propellant electrical isolator. Experimental results are presented indicating that all elements of the thruster design function well at the higher specific impulse and power levels. It is shown that the only thruster modifications required for extended-performance operation are a respacing of the ion optics assembly and a redesign of the propellant isolators. Experimental results obtained from three isolator designs are presented, and it is concluded that the design and development of a high-voltage isolator is possible using existing technology.

Beattie, J. R.↗

NASA NTRS1985

Scramjet sidewall burning: Preliminary shock tunnel results

Experiments performed with a two dimensional model scramjet with particular emphasis on the effect of fuel injection from a wall are reported. Air low with a nominal Mach number of 3.5 and varied enthalpies was produced. It was found that neither hydrogen injection angle nor combustor divergence angle had any appreciable effect on thrust values while increased combustor length appeared to increase thrust levels. Specific impulse was observed to peak when hydrogen was injected at an equivalence ratio of about 2. Lowering the Mach number of the injected hydrogen at low equivalence ratios, less than 4, appeared to benefit specific impulse while hydrogen Mach number had little effect at higher equivalence ratios. When a 1:1 mixture by volume of nitrogen and oxygen is used instead of air as a test gas, it is found that hydrogen combustion is enhanced but only at high enthalpies.

Morgan, R. G.↗

NASA NTRS2015

Propulsion System Testing for the Iodine Satellite (iSAT) Demonstration Mission

CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload, providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cm cu and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high delta v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. 3, 4 Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ρIsp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature (less than 100 C) to yield I2 vapor at or below 50 torr. At low power, the measured performance of an iodine-fed HET is very similar to that of a state-of-the-art xenon-fed thruster. Just as importantly, the current-voltage discharge characteristics of low power iodine-fed and xenon-fed thrusters are remarkably similar, potentially reducing development and qualifications costs by making it possible to use an already-qualified xenon-HET PPU in an iodine-fed system. Finally, a cold surface can be installed in a vacuum test chamber on which expended iodine propellant can deposit. In addition, the temperature doesn't have to be extremely cold to maintain a low vapor pressure in the vacuum chamber (it is under 10(exp -6) torr at −75 C), making it possible to 'cryopump' the propellant with lower-cost recirculating refrigerant-based systems as opposed to using liquid nitrogen or low temperature gaseous helium cryopanels. An iodine-based system is not without its challenges. The primary challenge is that the entire feed system must be maintained at an elevated temperature to prevent the iodine from depositing (transitioning from the gas phase directly back into the solid phase), which will block the propellant feed lines. Furthermore, deposition will occur unless the temperature in the lines is not greater than the temperature of the propellant reservoir. The flow rate can be controlled by adjusting the heating applied to the reservoir, but as with any thermal control there is a relatively slow response to changes in the heating rate. In the present paper, we describe the propulsion and propellant feed system for the iodine satellite (iSAT) flight demonstration mission. The system is based around the Busek BHT-200 Hall thruster, which has been modified for chemical compatibility with iodine vapor. While the gross propellant flow rate is maintained by the heated propellant reservoir, the flow to the anode and cathode are adjusted using two heated Vacco proportional flow control valves (PFCV), which provide very fast response on the flow rate adjustment. The flight mission design layout will be presented, showing how the system will be packaged into the overall 12-U spacecraft and the techniques being employed to protect the remaining spacecraft hardware from the propulsion system (e.g., plasma impingement, iodine deposition, thermal loads). In addition to the flight system design, results of testing the thruster and cathode with both operating on iodine propellant are presented. The tests are conducted on a thrust stand (see Fig. 1) in a large vacuum chamber containing a beam dump chilled to below -100 C to 'cryopump' the propellant. The thruster performance during these tests is presented, with these data used to evaluate the feed system and guide further refinements. Results of relatively long duration testing are presented to demonstrate the capability to operate for the length of the iSAT mission and to perform a number of re-starts as will be required by the mission concept of operations.

Polzin, Kurt A.↗

NASA NTRS2016

Engineering Model Propellant Feed System Development for an Iodine Hall Thruster Demonstration Mission

CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload, providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cu cm and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high (Delta)v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. 3, 4 Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ρIsp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature (< 100 C) to yield I2 vapor at or below 50 torr. At low power, the measured performance of an iodine-fed HET is very similar to that of a state-of-the-art xenon-fed thruster. Just as importantly, the current-voltage discharge characteristics of low power iodine-fed and xenon-fed thrusters are remarkably similar, potentially reducing development and qualifications costs by making it possible to use an already-qualified xenon-HET PPU in an iodine-fed system. Finally, a cold surface can be installed in a vacuum test chamber on which expended iodine propellant can deposit. In addition, the temperature doesn't have to be extremely cold to maintain a low vapor pressure in the vacuum chamber (it is under 10(exp −6) torr at −75 C), making it possible to 'cryopump' the propellant with lower-cost recirculating refrigerant-based systems as opposed to using liquid nitrogen or low temperature gaseous helium cryopanels. In the present paper, we describe the design and testing of the engineering model propellant feed system for iSAT (see Fig. 1). The feed system is based around an iodine propellant reservoir and two proportional control valves (PFCVs) that meter the iodine flow to the cathode and anode. The flow is split upstream of the PFCVs to both components can be fed from a common reservoir. Testing of the reservoir is reported to demonstrate that the design is capable of delivering the required propellant flow rates to operate the thruster. The tubing and reservoir are fabricated from hastelloy to resist corrosion by the heated gaseous iodine propellant. The reservoir, tubing, and PFCVs are heated to ensure the sublimed propellant will not re-deposit within the feed system. Heating is accomplished through a number of individual zones to control the overall power expended on heating the system and insulation is employed to minimize the amount of power used to heat the system prior to thruster operation.

Polzin, Kurt A.↗

NASA NTRS2017

Development of a 50,000-s, Lithium-fueled, Gridded Ion Thruster

The ion propulsion system on NASA’s Dawn mission provided over 11 km/s delta-V to the spacecraft. There is potential interest in missions that have delta-V’s an order of magnitude greater than this, i.e., 100 km/s to 200 km/s. To perform such missions would require a thruster that can produce a specific impulse roughly an order of magnitude greater than the 3100 s of the Dawn ion propulsion system. A 50-kW gridded ion thruster is being developed for operation with lithium propellant to produce a specific impulse of 50,000 s. The resulting thruster design requires a net accelerating voltage of 9 kV and a beam current of 5.5 A. Discharge chamber modeling is used to design a 35-cm diameter ring-cusp discharge chamber with six magnet rings. The discharge chamber is masked down to produce an active grid area that is ~25 cm diameter. Modeling suggests that the unique ionization characteristics of lithium may enable discharge chamber operation at a propellant efficiency of 99%. Operation at such a high propellant efficiency could significantly reduce the production of charge-exchange ions and thereby significantly reduce erosion of the accelerator grid.

Goebel, Dan M.↗

NASA NTRS1987

Experimental study of low Reynolds number nozzles

High-performance electrothermal thrusters operate in a low nozzle-throat Reynolds number regime. Under these conditions, the flow boundary layer occupies a large volume inside the nozzle, contributing to large viscous losses. Four nozzles (conical, bell, trumpet, and modified trumpet) and a sharp-edged orifice were evaluated over a Reynolds number range of 500 to 9000 with unheated nitrogen and hydrogen. The nozzles showed significant decreases in specific impulse efficiency with decreasing Reynolds number. At Reynolds numbers less than 1000, all four nozzles were probably filled with a large boundary layer. The discharge coefficient decreased with Reynolds number in the same manner as the specific impulse efficiency. The bell and modified trumpet nozzles had discharge coefficients 4 to 8 percent higher than those of the cone or trumpet nozzles. The Two-Dimensional Kinetics (TDK) nozzle analysis computer program was used to predict nozzle performance. The results were then compared to the experimental results in order to determine the accuracy of the program within this flow regime.

Grisnik, Stanley P.↗

NASA NTRS1987

Experimental study of low Reynolds number nozzles

Grisnik, Stanley P.↗

NASA NTRS1998

Scaling of Ion Thrusters to Low Power

Analyses were conducted to examine ion thruster scaling relationships in detail to determine performance limits, and lifetime expectations for thruster input power levels below 0.5 kW. This was motivated by mission analyses indicating the potential advantages of high performance, high specific impulse systems for small spacecraft. The design and development status of a 0.1-0.3 kW prototype small thruster and its components are discussed. Performance goals include thruster efficiencies on the order of 40% to 54% over a specific impulse range of 2000 to 3000 seconds, with a lifetime in excess of 8000 hours at full power. Thruster technologies required to achieve the performance and lifetime targets are identified.

Patterson, Michael J.↗

NASA NTRS2000

Advanced Hydrocarbon Fuel Development

As a part of a high energy density materials (HEDM) development, the hot fire tests for Quadricyclane, 1,7 Octadiyne, AFRL-1, Biclopropylidene, and CINCH (Dimethyl amino ethyl azide) have been conducted at NASA/MSFC. The first 4 materials for this task are provided from Air Force Research Laboratory at Edward Air Force Base and US Army provided CINCH. The performance of these fuels is compared with RP-1. The preliminary results of these tests are presented. The preliminary results of Quadricyclane tests indicate that the specific impulse and c-star efficiency for quadricyclane at the mixture ratio 1.94 are approximately 5 sec and 105 ft/sec better than the RP-1 at mixture ratio 1.9. The 1,7 Octadiyne test indicate that the specific impulse and c-star efficiency at the mixture ratio 2.1 are approximately -1 sec and 89 ft/sec differ than the RP-1 at mixture ratio 2.04. The Quadricyclane soot buildup at the combustor is a little more than RP-1, but detail study of soot formation is not considered at this time. There was no visual soot buildup for the 1,7 Octadiyne and AFRL-1.

Bai, S. Don↗

NASA NTRSDate not supplied

Effect of Engine Thrust and Isp Tradeoffs and Alternate Propellants on ΔV Budget and Architecture Mass for 1st Generation Nuclear Thermal Propulsion Flight Test Systems

Following the first Nuclear Thermal Propulsion (NTP) system test, also known as DRACO, the next NTP system to be developed would be the 1st Generation NTP system. An analysis was conducted to determine the performance of different vehicle configurations utilizing hydrogen (H-NTP), ammonia (A-NTP), and methane (M-NTP) as propellants launched onboard commercial launch vehicles. This was enabled by a quasi-steady-state power balance engine model and vehicle component physics that sized the vehicle system using Master Equipment List (MEL) parameters. The analysis considered configuration cases outlined by a matrix of different mission classes and vehicle configurations that covered the design space of the 1st Generation NTP system to explore various propellant options, engine architectures, and mission scopes. Four mission classes were evaluated which included single burn missions performing maneuvers having a ΔV of 1 km/s and 2.5 km/s followed by 2-burn and 4-burn missions that aimed to exhaust the launch vehicles volume or mass limitations. In all cases, the NTP systems with the lowest thrust class had the longest burn time of which A-NTP and M-NTP systems provided the longest and shortest burn times depending on the launch vehicle used while H-NTP systems tended to cluster together in the middle. Longer burn times could be useful as a testing platform to increase the time for data accumulation. Across the multi-burn cases, H-NTP systems were found to be volume limited while A-NTP and M-NTP systems were mass limited. Both single burn missions showed that A-NTP configurations provided the lowest dry mass given that ammonia had the highest density with comparable performance to M-NTP systems and no requirement for cryocoolers. The results showed that beyond the propellant selection type, the launch vehicle selection, which included Starship, New Glenn, Vulcan, Falcon 9 (recoverable), and Falcon Heavy (recoverable), was a primary driving factor in the test vehicle’s capabilities. Trends were determined based on a set of dimensionless parameters that included the mass ratio of inert mass to initial wetted mass, ratio of specific impulse to burn time, and a dimensionless engine parameter (ratio of impulse to engine momentum). These relationships found the “knees-in-the-curves” that could be a significant point of reference for the designer as they indicate a change in the trend which is located at a specific impulse to burn time ratio of 1, a mass ratio of 0.6, and an engine performance parameter of 4. This study did not attempt to make a recommendation rather provide the tools for the reader to select their own configuration based on their needs.

Propellant↗