Search NASA⌕ Search

SEARCH · Search NASA

Results for “Isp”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

Design and performance of an arcjet nuclear electric propulsion system for a mid-1990's reference mission

The design and performance of an arcjet nuclear-electric-propulsion spacecraft, suitable for use in the Space Nuclear Power System (SNPS) reference mission, are outlined. The vehicle design was based on a 30-kW ammonia arcjet system operating at an Isp of 1050 s and an efficiency of 45 percent. The arcjet/gimbal system, power-processing unit, and propellant-feed system are described. A 100-kWe SNPS was assumed, and the spacecraft mass was baselined at 5250 kg (excluding the propellant-feed system). A radiation/arcjet efflux diagnostics package was included in the performance analysis. This spacecraft, assuming a Shuttle launch from KSC, can perform a 50-deg inclination change and reach a final orbit of 35,860 km with a 120-d trip time providing a 4-mo active load for the SNPS. Alternatively, a Titan IV launch would provide a mass margin of 120 kg to a 10,000-km, 58-deg final orbit in 74 d. This spacecraft meets the reference-mission constraint of low developmental risk, and is scalable to power levels projected for future space platforms.

Deininger, William D.↗
NASA NTRS1988

A tandem mirror hybrid plume plasma propulsion facility

A concept in electrodeless plasma propulsion, which is also capable of delivering a variable Isp, is presented. The concept involves a three-stage system of plasma injection, heating, and subsequent ejection through a magnetic nozzle. The nozzle produces the hybrid plume by the coaxial injection of hypersonic neutral gas. The gas layer, thus formed, protects the material walls from the hot plasma and, through increased collisions, helps detach it from the diverging magnetic field. The physics of this concept is evaluated numerically through full spatial and temporal simulations; these explore the operating characteristics of such a device over a wide region of parameter space. An experimental facility to study the plasma dynamics in the hybrid plume was built. The device consists of a tandem mirror operating in an asymmetric mode. A later upgrade of this system will incorporate a cold plasma injector at one end of the machine. Initial experiments involve the full characterization of the operating envelope, as well as extensive measurements of plasma properties at the exhaust. The results of the numerical simulations are described.

Chang-Diaz, F. R.↗
NASA NTRS1989

Electric propulsion applications enabled by space nuclear power

Electric propulsion promises the advantage of providing high Isp's for placing payloads into their assigned orbits. This translates into heavier payloads using a given lift capability or, conversely, the use of smaller boosters. To accomplish this, high electric powers are required. Space reactor power systems such as SP-100 enable this technology. The electric propulsion requirements needed, namely, their power requirements and the resulting payload masses and time-to-orbit, are shown. Also indicated are the missions most benefitting from the use of electric propulsion. An Interim Reference Mission is described, synthesizing the results shown, for demonstration purposes.

Vicente, F. A.↗
NASA NTRS1990

Nuclear stage configuration studies for Mars missions

Several nuclear-propulsion stage configuration options for the February 2016 Mars Exploration Initiative mission are examined. Special attention is given to the 75,000-lb-thrust nuclear engine used for the major propulsive maneuvers. The reference mission, which will last 435 days, with the stay on Mars for 30 days, assumes that the nuclear thermal rocket (NTR) engine delivers a specific impulse of 925 sec with an engine thrust-to-weight ratio of 4. Results are given of trade studies performed on the NTR engine thrust level, engine thrust-to-weight ratio, and engine Isp. Attention is also given to the Mars transfer vehicle, the Mars transportation system (MTS), the MTS crew compartment, and the MTS tanks.

Emrich, W. J., Jr.↗
NASA NTRS1991

Uranium droplet core nuclear rocket

Uranium droplet nuclear rocket is conceptually designed to utilize the broad temperature range ofthe liquid phase of metallic uranium in droplet configuration which maximizes the energy transfer area per unit fuel volume. In a baseline system dissociated hydrogen at 100 bar is heated to 6000 K, providing 2000 second of Isp. Fission fragments and intense radian field enhance the dissociation of molecular hydrogen beyond the equilibrium thermodynamic level. Uranium droplets in the core are confined and separated by an axisymmetric vortex flow generated by high velocity tangential injection of hydrogen in the mid-core regions. Droplet uranium flow to the core is controlled and adjusted by a twin flow nozzle injection system.

Anghaie, Samim↗
NASA NTRS1991

Nuclear rocket using indigenous Martian fuel NIMF

In the 1960's, Nuclear Thermal Rocket (NTR) engines were developed and ground tested capable of yielding isp of up to 900 s at thrusts up to 250 klb. Numerous trade studies have shown that such traditional hydrogen fueled NTR engines can reduce the inertial mass low earth orbit (IMLEO) of lunar missions by 35 percent and Mars missions by 50 to 65 percent. The same personnel and facilities used to revive the hydrogen NTR can also be used to develop NTR engines capable of using indigenous Martian volatiles as propellant. By putting this capacity of the NTR to work in a Mars descent/acent vehicle, the Nuclear rocket using Indigenous Martian Fuel (NIMF) can greatly reduce the IMLEO of a manned Mars mission, while giving the mission unlimited planetwide mobility.

Zubrin, Robert↗
NASA NTRS1991

Droplet Core Nuclear Rocket (DCNR)

The most basic design feature of the droplet core nuclear reactor is to spray liquid uranium into the core in the form of droplets on the order of five to ten microns in size, to bring the reactor to critical conditions. The liquid uranium fuel ejector is driven by hydrogen, and more hydrogen is injected from the side of the reactor to about one and a half meters from the top. High temperature hydrogen is expanded through a nozzle to produce thrust. The hydrogen pressure in the system can be somewhere between 50 and 500 atmospheres; the higher pressure is more desirable. In the lower core region, hydrogen is tangentially injected to serve two purposes: (1) to provide a swirling flow to protect the wall from impingement of hot uranium droplets: (2) to generate a vortex flow that can be used for fuel separation. The reactor is designed to maximize the energy generation in the upper region of the core. The system can result in and Isp of 2000 per second, and a thrust-to-weight ratio of 1.6 for the shielded reactor. The nuclear engine system can reduce the Mars mission duration to less than 200 days. It can reduce the hydrogen consumption by a factor of 2 to 3, which reduces the hydrogen load by about 130 to 150 metric tons.

Anghaie, Samim↗
NASA NTRS1991

Lunar mission aerobrake performance study

Nine lunar mission scenarios were developed to show the transfer vehicle performance benefits of aerobraking into low Earth orbit (LEO) upon Earth return as opposed to an all-propulsive maneuver. The initial mass in LEO (IMLEO) of the lunar transfer vehicle is considered the measure of vehicle performance. Four types of mission profiles in conjunction with two vehicle concepts were used to construct the scenarios. These nine scenarios were designed to represent a broad range of possible lunar missions so that a general knowledge base of aerobraking and lunar transfer vehicle performance levels could be obtained. Also discussed are the mass sensitivities of each transfer vehicle to changes in the selected design parameters: Isp, crew module mass, payload to surface, and aerobrake mass fraction. A parametric study was performed on two of the mission scenarios to help quantify the performance benefits by adding a set of drop tanks to the vehicle. The parametric study also provides partial derivatives which show the sensitivities of IMLEO to the four design parameters listed. A ranking is given based on vehicle performance.

Mulqueen, J.↗
NASA NTRS1992

CFD analysis of the STME nozzle flowfield

The Space Transportation Main Engine (STME) uses a gas generator cycle to cool the nozzle wall by a film-dump of the turbine exhaust. The ability to cool the skirt is a key concern in the design of the STME. CFD calculations were undertaken to predict the film cooling effectiveness and performance sensitivities for various configurations, operating points, and inlet conditions. The results presented here were obtained for the subscale nozzle. The computations were performed using REFLEQS. The computational analysis showed that a chemical equilibrium model was necessary to obtain correct predictions of the specific impulse. The frozen composition model underpredicts the ISP by about 6 percent. It was also observed that the coolant film was successful in maintaining the nozzle wall well below the stagnation temperature of the core flow. The effect of the coolant flow on the performance of the engine was found to be negligible. The computed heat fluxes at the wall were in good agreement with the empirical data obtained by Pratt and Whitney. Further test data from Pratt and Whitney are forthcoming for the subscale nozzle. Calculations will be performed to determine cooling efficiencies and nozzle performance over a range of conditions, and model predictions will be compared with experimental data. Information is given in viewgraph form.

Krishnan, Anantha↗
NASA NTRS1993

Innovative electric propulsion thruster modeling

The objective of this program is to model and evaluate advanced nuclear electric propulsion (NEP) system concepts as an aid to the performance of NEP mission benefits studies. The two primary goals are as follows: (1) provide scaling relationships for mass, power, and efficiency, as functions of Isp, propellant type, and other important quantities. The discussion is presented in vugraph form.

Frisbee, Robert H.↗
NASA NTRS1993

An analytical optimization of electric propulsion orbit transfer vehicles

Due to the electric propulsion's inherent propellant mass savings over chemical propulsion, electric propulsion orbit transfer vehicles (EPOTV's) are highly efficient mode of orbit transfer. When selecting an electric propulsion device (ion, MPD, or arcjet) and propellant for a particular mission, it is preferable to use quick, analytical system optimization methods instead of time intensive numerical integration methods. It is also of interest to determine each thruster's optimal operating characteristics for a specific mission. Analytical expressions are derived which determine the optimal specific impulse (Isp) for each type of electric thruster to maximize payload fraction for a desired thrusting time. These expressions take into account the variation of thruster efficiency with specific impulse. Verification of the method is made with representative electric propulsion values on a LEO-to-GEO mission. Application of the method to specific missions is discussed.

Oleson, Steven R.↗
NASA NTRS1990

Propulsion needs for lunar/Mars missions

Viewgraphs on propulsion needs for lunar/Mars missions are presented. Topics covered include: lunar mission profile; lunar transfer vehicle/lunar excursion vehicle; launch vehicles for lunar missions; lunar outpost; Mars transfer operations; Mars mission vehicle in low earth orbit (LEO); mission vehicle commonality; mass comparison for reference missions; advanced propulsion; propulsion option size comparison; Mars transfer vehicle (MTV) propulsion option weights for mission favorable opportunities; propulsion options comparison; Mars transportation architecture options; tanker options for fully reusable systems; NTR 900 Isp staged tanks and engines; fully reusable cryogenic aerobraked system; NEP operated from high orbit; and nuclear safe orbit considerations.

Woodcock, Gordon R.↗
NASA NTRS1994

Fault tolerant onboard packet switch architecture for communication satellites: Shared memory per beam approach

The NASA Lewis Research Center is developing a multichannel communication signal processing satellite (MCSPS) system which will provide low data rate, direct to user, commercial communications services. The focus of current space segment developments is a flexible, high-throughput, fault tolerant onboard information switching processor. This information switching processor (ISP) is a destination-directed packet switch which performs both space and time switching to route user information among numerous user ground terminals. Through both industry study contracts and in-house investigations, several packet switching architectures were examined. A contention-free approach, the shared memory per beam architecture, was selected for implementation. The shared memory per beam architecture, fault tolerance insertion, implementation, and demonstration plans are described.

Shalkhauser, Mary JO↗
NASA NTRS1993

A minimum cost tolerance allocation method for rocket engines and robust rocket engine design

Rocket engine design follows three phases: systems design, parameter design, and tolerance design. Systems design and parameter design are most effectively conducted in a concurrent engineering (CE) environment that utilize methods such as Quality Function Deployment and Taguchi methods. However, tolerance allocation remains an art driven by experience, handbooks, and rules of thumb. It was desirable to develop and optimization approach to tolerancing. The case study engine was the STME gas generator cycle. The design of the major components had been completed and the functional relationship between the component tolerances and system performance had been computed using the Generic Power Balance model. The system performance nominals (thrust, MR, and Isp) and tolerances were already specified, as were an initial set of component tolerances. However, the question was whether there existed an optimal combination of tolerances that would result in the minimum cost without any degradation in system performance.

Gerth, Richard J.↗
NASA NTRS1995

An Accelerated Development, Reduced Cost Approach to Lunar/Mars Exploration Using a Modular NTR-Based Space Transportation System

The results of integrated systems and mission studies are presented which quantify the benefits and rationale for developing a common, modular lunar/Mars space transportation system (STS) based on nuclear thermal rocket (NTR) technology. At present NASA's Exploration Program Office (ExPO) is considering chemical propulsion for an 'early return to the Moon' and NTR propulsion for the more demanding Mars missions to follow. The time and cost to develop these multiple systems are expected to be significant. The Nuclear Propulsion Office (NPO) has examined a variety of lunar and Mars missions and heavy lift launch vehicle (HLLV) options in an effort to determine a 'standardized' set of engine and stage components capable of satisfying a wide range of Space Exploration Initiative (SEI) missions. By using these components in a 'building block' fashion, a variety of single and multi-engine lunar and Mars vehicles can be configured. For NASA's 'First Lunar Outpost' (FLO) mission, an expendable NTR stage powered by two 50 klbf engines can deliver approximately 96 metric tons (t) to translunar injection (TLI) conditions for an initial mass in low earth orbit (IMLEO) of approximately 198 t compared to 250 t for a cryogenic chemical TLI stage. The NTR stage liquid hydrogen (LH2) tank has a 10 m diameter, 14.5 m length, and 66 t LH2 capacity. The NTR utilizes a UC-ZrC-graphite 'composite' fuel with a specific impulse (Isp) capability of approximately 900 s and an engine thrust-to-weight ratio of approximately 4.3. By extending the size and LH2 capacity of the lunar NTR stage to approximately 20 m and 96 t, respectively, a single launch Mars cargo vehicle capable of delivering approximately 50 t of surface payload is possible. Three 50 klbf NTR engines and the two standardized LH2 tank sizes developed for lunar and Mars cargo vehicle applications would be used to configure the Mars piloted vehicle for a mission as early as 2010. The paper describes the features of the 'common' NTR-based moon/Mars STS, examines performance sensitivities resulting from different 'mission mode' assumptions, and quantifies potential schedule and cost benefits resulting from this modular moon/Mars NTR vehicle approach.

Borowski, S.↗
NASA NTRS1995

Design Considerations for Space Transfer Vehicles Using Solar Thermal Propulsion

The economical deployment of satellites to high energy earth orbits is crucial to the ultimate success of this nations commerical space ventures and is highly desirable for deep space planetary missions requiring earth escape trajectories. Upper stage space transfer vehicles needed to accomplish this task should ideally be simple, robust, and highly efficient. In this regard, solar thermal propulsion is particularly well suited to those missions where high thrust is not a requirement. The Marshall Space Flight Center is , therefore, currently engaged in defining a transfer vehicle employing solar thermal propulsion capable of transferring a 1000 lb. payload from low Earth orbit (LEO) to a geostationary Earth orbit (GEO) using a Lockheed launch vehicle (LLV3) with three Castors and a large shroud. The current design uses liquid hydrogen as the propellant and employs two inflatable 16 x 24 feet eliptical off-axis parabolic solar collectors to focus sunlight onto a tungsten/rhenium windowless black body type absorber. The concentration factor on this design is projected to be approximately 1800:1 for the primary collector and 2.42:1 for the secondary collector for an overall concentration factor of nearly 4400:1. The engine, which is about twice as efficient as the best currently available chemical engines, produces two pounds of thrust with a specific impulse (Isp) of 860 sec. Transfer times to GEO are projected to be on the order of one month. The launch and deployed configurations of the solar thermal upper stage (STUS) are depicted.

Emrich, William J.↗
NASA NTRS1995

Influence of Power System Technology on Electric Propulsion Missions

Electric propulsion (EP) thruster technology, with efficient lightweight power systems can provide substantial reductions in propulsion system wet mass due to the high specific impulse (Isp) of the thrusters. Historically, the space power systems are too massive for many potential orbital missions. The objective of this paper is to show the impact of current power system technology on EP mission performance and determine what technology advancements are needed to make EP beneficial for earth orbital applications. The approach of the paper is to model the electric propulsion system and orbital mission using a partial parametric method. Various missions are analyzed from orbit maintenance to orbit transfer. Results portray the relationship between mission performance and power technology level. Conclusions show which mission applications currently have acceptable power technology, and which mission applications require power technology improvements.

Oleson, Steven R.↗
NASA NTRS1997

High Pressure, Earth-storable Rocket Technology

The effect of elevated chamber pressure on combustion efficiency and heat transfer has been determined at the 100 lbf (445 N) thrust level for nitrogen tetroxide propellants. Measurements were made up to 500 psia (3.45 Mpa) with testbed hardware; tests at 100 psia (0.690 MPa) and 250 psia (1.72 MPa) were made with radiation-cooled rhenium chambers. The first task of the program served to determine desirable thruster applications and operating conditions: high total impulse, i.e. communication satellite or spacecraft bus axial engines, at chamber pressures up to 250 psia (1.72 MPa) pressure-fed, or up to 500 psia (3.45 MPa) pump-fed. The hardware modifications and testing required to obtain the data were determined in Task 2, which included design-support hot fire tests; supplemental hardware, including a 250 psia (1.72 MPa) Pc rhenium chamber and a 20% fuel-film cooled platelet injector was fabricated in Task 3. Testing showed that satisfactory operation of Ir-Re radiation chambers is assured at pressures up to 250 psia and may be possible up to 500. The heat transfer data obtained show good correlation with throat Reynolds number and are generally under values given by the simplified Bartz equation; chambers equilibrium temperatures match predicted values. Preliminary optimization of trip configuration and mixture ratio were made; Isp performance from thrust measurements was within 1% of predicted values. Stability, compatibility, and front-end thermal management were determined to be satisfactory.

Jassowski, D. M.↗