Search NASA⌕ Search

SEARCH · Search NASA

Results for “RP”

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 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

Chondrules in the Qingzhen type-3 enstatite chondrite Possible precursor components and comparison to ordinary chondrite chondrules

The mineral composition of chondrules from a fragment of Qingzhen (EH3) fall was analyzed by neutron activation method. Unlike the ordinary chondrite (OC) chondrules (Gooding and Keil, 1981), the Qingzhen radial pyroxene (RP) and porphyritic pyroxene (PP) chondrules have similar bulk compositions. Porphyritic olivine-pyroxene (POP) chondrules are richer than PP and RP chondrules in refractory lithophiles and siderophiles. Elements in each of the following sets intercorrelate significantly: (1) Fe-Co-Ni-Ir-Au, probably derived from a metal component; (2) Ca-Eu-Se, which suggests an oldhamite-rich precursor; (3) Al-Sc-Hf, occurring in high concentrations in POP chondrules, this set suggesting the existence of a refractory lithophile-rich and olivine-rich component; (4) Na REE; and (5) Cl-Br. Sets (2) and (4) were not precursors of OC. The interelement ratios of refractory lithophiles such as Ca, Al, Ti, Sc, and REE are similar to CI ratios, suggesting that they originated in the earliest phases as silicates, which were sulfurized before chondrule formation.

Grossman, J. N.↗
NASA NTRS1989

Liquid Oxygen Cooling of Hydrocarbon Fueled Rocket Thrust Chambers

Rocket engines using liquid oxygen (LOX) and hydrocarbon fuel as the propellants are being given serious consideration for future launch vehicle propulsion. Normally, the fuel is used to regeneratively cool the combustion chamber. However, hydrocarbons such as RP-1 are limited in their cooling capability. Another possibility for the coolant is the liquid oxygen. Combustion chambers previously tested with LOX and RP-1 as propellants and LOX as the collant demonstrated the feasibility of using liquid oxygen as a coolant up to a chamber pressure of 13.8 MPa (2000 psia). However, there was concern as to the effect on the integrity of the chamber liner if oxygen leaks into the combustion zone through fatigue cracks that may develop between the cooling passages and the hot gas side wall. In order to study this effect, chambers were fabricated with slots machined upstream of the throat between the cooling passage wall and the hot gas side wall to simulate cracks. The chambers were tested at a nominal chamber pressure of 8.6 MPa (1247 psia) over a range of mixture ratios from 1.9 to 3.1 using liquid oxygen as the coolant. The results of the testing showed that the leaking LOX did not have a deleterious effect on the chambers in the region of the slots. However, there was unexplained melting in the throat region of both chambers, but not in line with the slots.

Armstrong, Elizabeth S.↗
NASA NTRS1989

Liquid oxygen cooling of hydrocarbon fueled rocket thrust chambers

Rocket engines using liquid oxygen (LOX) and hydrocarbon fuel as the propellants are being given serious consideration for future launch vehicle propulsion. Normally, the fuel is used to regeneratively cool the combustion chamber. However, hydrocarbons such as RP-1 are limited in their cooling capability. Another possibility for the coolant is the liquid oxygen. Combustion chambers previously tested with LOX and RP-1 as propellants and LOX as the coolant demonstrated the feasibility of using liquid oxygen as a coolant up to a chamber pressure of 13.8 MPa (2000 psia). However, there was concern as to the effect on the integrity of the chamber liner if oxygen leaks into the combustion zone through fatigue cracks that may develop between the cooling passages and the hot gas side wall. In order to study this effect, chambers were fabricated with slots machined upstream of the throat between the cooling passage wall and the hot gas side wall to simulate cracks. The chambers were tested at a nominal chamber pressure of 8.6 MPa (1247 psia) over a range of mixture ratios from 1.9 to 3.1 using liquid oxygen as the coolant. The results of the testing showed that the leaking LOX did not have a deleterious effect on the chambers in the region of the slots. However, there was unexplained melting in the throat region of both chambers, but not in line with the slots.

Armstrong, Elizabeth S.↗
NASA NTRS1989

Atlas of albedo and absorbed solar radiation derived from Nimbus 6 earth radiation budget data set, July 1975 to May 1978

An atlas of monthly mean global contour maps of albedo and absorbed solar radiation is presented. The atlas is based on 35 months of continuous measurements from July 1975 through May 1978. The data were retrieved from measurements made by the shortwave wide field-of-view radiometer of the first Earth Radiation Budget (ERB) instrument, which flew on the Nimbus 6 spacecraft in 1975. Profiles of zonal mean albedos and absorbed solar radiation are tabulated. These geographical distributions are provided as a resource for studying the radiation budget of the earth. This atlas of albedo and absorbed solar radiation complements the atlases of outgoing longwave radiation by Bess and Smith in NASA-RP-1185 and RP-1186, also based on the Nimbus 6 and 7 ERB data.

Smith, G. Louis↗
NASA NTRS1989

The measurement, modeling, and prediction of traction for rocket propellant 1

Traction tests were performed on RP-1, a common kerosene based rocket propellant. Traction data on this fluid are required for purposes of turbopump bearing design, using codes such as SHABERTH. To obtain the traction data, an existing twin disc machine was used, operating under the side slip mode and using elliptical contacts. The range of test variables were: contact peak Hertz stress from 1.0 to 2.0 GPa, disc surface speed from 10 to 50 m/s, fluid inlet temperature from 30 to 70 C, and with a contact aspect ratio of 1.7. The resulting traction curves were reduced to fundamental fluid property parameters using the Johnson and Tevaarwerk traction model. Theoretical traction predictions were performed by back substitution of the fundamental properties into the traction model. Comparison of the predicted with the measured curves gives a high degree of confidence in the correctness of the traction model. For purposes of input to the NASA SHABERTH program, the traction model was next used to predict the expected traction of RP-1 under line contact conditions.

Tevaarwerk, J. L.↗
NASA NTRS1990

Cooling of rocket thrust chambers with liquid oxygen

Rocket engines using high pressure liquid oxygen (LOX) and kerosene (RP-1) as the propellants have been considered for future launch vehicle propulsion. Generally, in regeneratively cooled engines, the fuel is used to cool the combustion chamber. However, hydrocarbons such as RP-1 are limited in their cooling capability at high temperatures and pressures. Therefore, LOX is being considered as an alternative coolant. However, there has been concern as to the effect on the integrity of the chamber liner if oxygen leaks into the combustion zone through fatigue cracks that may develop between the cooling passages and the hot-gas side wall. To address this concern, an investigation was previously conducted with simulated fatigue cracks upstream of the thrust chamber throat. When these chambers were tested, an unexpected melting in the throat region developed which was not in line with the simulated fatigue cracks. The current experimental program was conducted in order to determine the cause for the failure in the earlier thrust chambers and to further investigate the effects of cracks in the thrust chamber liner upstream of the throat. The thrust chambers were tested at oxygen-to-fuel mixture ratios from 1.5 to 2.86 at a nominal chamber pressure of 8.6 MPa. As a result of the test series, the reason for the failure occurring in the earlier work was determined to be injector anomalies. The LOX leaking through the simulated fatigue cracks did not affect the integrity of the chambers.

Armstrong, Elizabeth S.↗
NASA NTRS1989

Size-frequency distributions of chondrules in CO3 chondrites

The size-frequency distributions of chondrules were determined for eleven CO3 chondrites (including ALHA77003, ALHA77307, ALH82101, ALH85003, Colony, Felix, Isna, Kainsaz, Lance, Ornans, and Warrenton), using the results of petrographic analyses of thin sections. The mean proportion of different chondrule types in CO3 chondrites were estimated to be 69 percent POP, 18 percent PP, 8 percent PO, 2 percent BO, 2 percent RP, 1 percent C, and less than 0.1 percent GOP. These proportions are very different from those in ordinary or EH and CV chondrites, with the smaller proportion of nonporphyritic chondrules than EH chondrites, but a larger proportion than CV chondrites. Relative proportions of chondrule types vary with size interval; thus, with decreasing chondrule size, PO chondrules decrease fairly regularly in abundance, while RP chondrules are most abundant in the smallest size intervals.

Rubin, Alan E.↗
NASA NTRS1990

Microwave dielectric properties of dry rocks

A combination of techniques was used to measure the dielectric properties of 80 rock samples in the microwave region. The real part (RP) of the relative dielectric constant was measured in 0.1-GHz steps from 0.5 to 18 GHz, and the imaginary part (IP) was measured at five frequencies between 1.6 and 16 GHz. The bulk density rho(b) was also measured for all the samples, and the bulk chemical composition (BCC) was determined for 56 of the samples. RP is found to be frequency-independent at 0.5-18 GHz for all samples, and rho(b) accounts for about 50 percent of the observed variance. For silicate rocks, as much as 78 percent of the variance is explained by the combination of rho(b) and the fractional contents of oxides when the silicates are subgrouped by genesis. In contrast, IP decreases with increasing frequency for most rock samples, and no statistically significant relationships are found between IP and rho(b). For subgrouped silicate rocks, 60 percent of the variance in IP can be explained by BCC.

Ulaby, Fawwaz T.↗
NASA NTRS1990

Cooling of rocket thrust chambers with liquid oxygen

Rocket engines using high pressure liquid oxygen (LOX) and kerosene (RP-1) as the propellants have been considered for future launch vehicle propulsion. Generaly, in regeneratively cooled engines, thefuel is used to cool the combustion chamber. However, hydrocarbons such as RP-1 are limited in their cooling capability at high temperatures and pressures. Therefore, LOX is being considered as an alternative coolant. However, there has been concern as to the effect on the integrity of the chamber liner if oxygen leaks into the combustion zone through fatigue cracks that may develop between the cooling passages and the hot-gas side wall. To address this concern, an investigation was previously conducted with simulated fatigue cracks upstream of the thrust chamber throat. When these chambers were tested, an unexpected melting in the throat region developed which was not in line with the simulated fatigue cracks. The current experimental program was conducted in order to determine the cause for the failure in the earlier thrust chambers and to further investigate the effects of cracks in the thrust chamber liner upstream of the throat. The thrust chambers were tested at oxygen-to-fuel mixture ratios from 1.5 to 2.86 at a nominal chamber pressure of 8.6 MPa. As a result of the test series, the reason for the failure occurring in the earlier work was determined to be injector anomalies. The LOX leaking through the simulated fatigue cracks did not affect the integrity of the chambers.

Armstrong, Elizabeth S.↗
NASA NTRS1990

A synopsis of Langley Research Center's lidar effort for the 1986 FIRE IFO

The lidar data obtained by the Langley Aircraft Lidar in October 1986 in Wisconsin is being reduced in a transparent, simple fashion and will be published in its reduced form in a NASA Reference Publication (RP). This reduced data will also be submitted to the FIRE data archives. Some of this reduced data will be presented at the FIRE FSET Workshop to acquaint the science team with the data format to be used in the archive and the upcoming catalog contained in the RP. A new method was utilized in Wisconsin for obtain the depolarization ratio of aerosols. This method involves using a half-wave plate to calibrate the lidar under field conditions. The theory behind this technique will be presented at this workshop as well as some of the lidar calibration results. The lidar calibration will be utilized in interpreting some of the dual polarization lidar data obtained during the IFO in Wisconsin. Some of these data are also discussed. A continuous wave laser lab-type lidar simulator was constructed during the previous year. One of the primary reasons for the construction of the simulator was to attempt dual-polarization lidar-like calibrations under laboratory, rather than field conditions. The data collected by this system was used to experimentally check and thus, inspire confidence in the algorithms being used to interpret the lidar data obtained in the field. A computer program which simulates noisy lidar data was used as a part of this effort in order to obtain some feel for the noise in the inversion parameters as a function of noise in the actual measurements. The lidar simulation will be described in addition to presenting some of the lab-generated calibration data.

Alvarez, Jose M.↗
NASA NTRS1991

An Explicit Upwind Algorithm for Solving the Parabolized Navier-Stokes Equations

An explicit, upwind algorithm was developed for the direct (noniterative) integration of the 3-D Parabolized Navier-Stokes (PNS) equations in a generalized coordinate system. The new algorithm uses upwind approximations of the numerical fluxes for the pressure and convection terms obtained by combining flux difference splittings (FDS) formed from the solution of an approximate Riemann (RP). The approximate RP is solved using an extension of the method developed by Roe for steady supersonic flow of an ideal gas. Roe's method is extended for use with the 3-D PNS equations expressed in generalized coordinates and to include Vigneron's technique of splitting the streamwise pressure gradient. The difficulty associated with applying Roe's scheme in the subsonic region is overcome. The second-order upwind differencing of the flux derivatives are obtained by adding FDS to either an original forward or backward differencing of the flux derivative. This approach is used to modify an explicit MacCormack differencing scheme into an upwind differencing scheme. The second order upwind flux approximations, applied with flux limiters, provide a method for numerically capturing shocks without the need for additional artificial damping terms which require adjustment by the user. In addition, a cubic equation is derived for determining Vegneron's pressure splitting coefficient using the updated streamwise flux vector. Decoding the streamwise flux vector with the updated value of Vigneron's pressure splitting improves the stability of the scheme. The new algorithm is applied to 2-D and 3-D supersonic and hypersonic laminar flow test cases. Results are presented for the experimental studies of Holden and of Tracy. In addition, a flow field solution is presented for a generic hypersonic aircraft at a Mach number of 24.5 and angle of attack of 1 degree. The computed results compare well to both experimental data and numerical results from other algorithms. Computational times required for the upwind PNS code are approximately equal to an explicit PNS MacCormack's code and existing implicit PNS solvers.

Korte, John J.↗
NASA NTRS1993

Multilayer thin film design for far ultraviolet polarizers using an induced transmission and absorption technique

An explanation of induced transmission for spectral regions excluding the far ultraviolet (FUV) is given to better understand how induced transmission and absorption can be used to design effective polarizers in the FUV spectral region. We achieve high s-polarization reflectance and a high degree of polarization (P equals (Rs-Rp)/(Rs+Rp)) by means of a MgF2/Al/MgF2 three layer structure on an opaque thick film of Al as the substrate. For example, our polarizer designed for the Lyman-alpha line (lambda equals 121.6 nm) has 87.95 percent reflectance for the s-polarization case and 0.43 percent for the p-polarization case, with a degree of polarization of 99.03 percent. If a double reflection polarizer is made with this design, it will have a degree of polarization of 99.99 percent and s-polarization throughput of 77.35 percent.

Kim, Jongmin↗
NASA NTRS1994

LSENS, a general chemical kinetics and sensitivity analysis code for homogeneous gas-phase reactions. 2: Code description and usage

LSENS, the Lewis General Chemical Kinetics Analysis Code, has been developed for solving complex, homogeneous, gas-phase chemical kinetics problems and contains sensitivity analysis for a variety of problems, including nonisothermal situations. This report is part 2 of a series of three reference publications that describe LSENS, provide a detailed guide to its usage, and present many example problems. Part 2 describes the code, how to modify it, and its usage, including preparation of the problem data file required to execute LSENS. Code usage is illustrated by several example problems, which further explain preparation of the problem data file and show how to obtain desired accuracy in the computed results. LSENS is a flexible, convenient, accurate, and efficient solver for chemical reaction problems such as static system; steady, one-dimensional, inviscid flow; reaction behind incident shock wave, including boundary layer correction; and perfectly stirred (highly backmixed) reactor. In addition, the chemical equilibrium state can be computed for the following assigned states: temperature and pressure, enthalpy and pressure, temperature and volume, and internal energy and volume. For static problems the code computes the sensitivity coefficients of the dependent variables and their temporal derivatives with respect to the initial values of the dependent variables and/or the three rate coefficient parameters of the chemical reactions. Part 1 (NASA RP-1328) derives the governing equations describes the numerical solution procedures for the types of problems that can be solved by lSENS. Part 3 (NASA RP-1330) explains the kinetics and kinetics-plus-sensitivity-analysis problems supplied with LSENS and presents sample results.

Radhakrishnan, Krishnan↗
NASA NTRS1996

Preliminary work toward the development of a dimensional tolerance standard for rapid prototyping

Rapid prototyping is a new technology for building parts quickly from CAD models. It works by slicing a CAD model into layers, then by building a model of the part one layer at a time. Since most parts can be sliced, most parts can be modeled using rapid prototyping. The layers themselves are created in a number of different ways - by using a laser to cure a layer of an epoxy or a resin, by depositing a layer of plastic or wax upon a surface, by using a laser to sinter a layer of powder, or by using a laser to cut a layer of paper. Rapid prototyping (RP) is new, and a standard part for use in comparing dimensional tolerances has not yet been chosen and accepted by ASTM (the American Society for Testing Materials). Such a part is needed when RP is used to build parts for investment casting or for direct use. The objective of this project was to start the development of a standard part by using statistical techniques to choose the features of the part which show curl - the vertical deviation of a part from its intended horizontal plane.

Kennedy, W. J.↗
NASA NTRS1998

X-34 Main Propulsion System Design and Operation

The X-34 program is a joint industry/government program to develop, test, and operate a small, fully-reusable hypersonic flight vehicle, utilizing technologies and operating concepts applicable to future Reusable Launch Vehicle (RLV) systems. The vehicle will be capable of Mach 8 flight to 250,000 feet altitude and will demonstrate an all composite structure, composite RP-1 tank, the Marshall Space Flight Center (MSFC) developed Fastrac engine, and the operability of an advanced thermal protection systems. The vehicle will also be capable of carrying flight experiments. MSFC is supporting the X-34 program in three ways: Program Management, the Fastrac engine as Government Furnished Equipment (GFE), and the design of the Main Propulsion System (MPS). The MPS Product Development Team (PDT) at MSFC is responsible for supplying the MPS design, analysis, and drawings to Orbital. The MPS consists of the LOX and RP-1 Fill, Drain, Feed, Vent, & Dump systems and the Helium & Nitrogen Purge, Pressurization, and Pneumatics systems. The Reaction Control System (RCS) design was done by Orbital. Orbital is the prime contractor and has responsibility for integration, procurement, and construction of all subsystems. The paper also discusses the design, operation, management, requirements, trades studies, schedule, and lessons learning with the MPS and RCS designs.

Champion, R. J., Jr.↗
NASA NTRS1998

Low Cost Propulsion Technology at the Marshall Space Flight Center: Fastrac Engine and the Propulsion Test Article

The need for low cost access to space has initiated the development of low cost liquid rocket engine and propulsion system hardware at the Marshall Space Flight Center (MSFC). The engine, the 60,000 lbf, RP-1 and LOX Fastrac Engine has been designed as a robust, low cost liquid rocket engine with applications for X-34 as well as future low cost booster systems. The engine is a turbopump fed, gas generator cycle, rocket motor with an ablative nozzle. The Propulsion Test Article (PTA) is a test bed for low cost propulsion system hardware including a composite RP-1 tank, flight feedlines and pressurization system, stacked in a booster configuration. A general description of the PTA and the Fastrac engine is given, with emphasis on the technical specification of the hardware including flow rates, pressures and other operating conditions. The process which has been used for the design and integration of this hardware is described.

Fisher, Mark F.↗
NASA NTRS1998

Low Cost Propulsion Technology Testing at the Stennis Space Center: Propulsion Test Article and the Horizontal Test Facility

The need for low cost access to space has initiated the development of low cost liquid rocket engine and propulsion system hardware at the Marshall Space Flight Center. This hardware will be tested at the Stennis Space Center's B-2 test stand. This stand has been reactivated for the testing of the Marshall designed Fastrac engine and the Propulsion Test Article. The RP-1 and LOX engine is a turbopump fed gas generator rocket with an ablative nozzle which has a thrust of 60,000 lbf. The Propulsion Test Article (PTA) is a test bed for low cost propulsion system hardware including a composite RP-I tank, flight feedlines and pressurization system, stacked in a booster configuration. The PTA is located near the center line of the B-2 test stand, firing vertically into the water cooled flame deflector. A new second position on the B-2 test stand has been designed and built for the horizontal testing of the Fastrac engine in direct support of the X-34 launch vehicle. The design and integration of these test facilities as well as the coordination which was required between the two Centers is described and lessons learned are provided. The construction of the horizontal test position is discussed in detail. The activation of these facilities is examined and the major test milestones are described.

Fisher, Mark F.↗
NASA NTRS1998

Hydrodynamic Design of the Fastrac Turbopump

A 60,000 lb trust liquid Oxygen (LOX) and rocket propellant #1 (RP-1) rocket engine is being developed at the Marshall Space Flight Center (MSFC) as part of NASA's ongoing effort to lower the cost of access to space. The goal of the program is to demonstrate a low cost rocket engine with acceptable performance. In order to reduce design and development risk, computational fluid dynamics (CFD) analysis was used extensively in the design process, especially in the design of the turbopump. This paper documents the process used and significant results obtained during the design of the LOX and RP-1 pumps. This includes the design of the inlet manifolds, the integral inducer-impellers, the diffusers, and the collector volutes. The resulting designs incorporate low-cost design features while meeting the engine cycle requirements.

Garcia, Roberto↗