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Lawver, B. R.

Publications and source records attributed to Lawver, B. R..

16 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

Ignition characterization of LOX/hydrocarbon propellants

The results of an evaluation of the ignition characteristics of the gaseous oxygen (Gox)/Ethanol propellant combination are presented. Ignition characterization was accomplished through the analysis, design, fabrication and testing of a spark initiated torch igniter and prototype 620 lbF thruster/igniter assembly. The igniter was tested over a chamber pressure range of 74 to 197 psia and mixture ratio range of 0.778 to 3.29. Cold (-92 to -165 F) and ambient (44 to 80 F) propellant temperatures were used. Spark igniter ignition limits and thruster steady state and pulse mode, performance, cooling and stability data are presented. Spark igniter ignition limits are presented in terms of cold flow pressure, ignition chamber diameter and mixture ratio. Thruster performance is presented in terms of vacuum specific impulse versus engine mixture ratio. Gox/Ethanol propellants were shown to be ignitable over a wide range of mixture ratios. Cold propellants were shown to have a minor effect on igniter ignition limits. Thruster pulse mode capability was demonstrated with multiple pulses of 0.08 sec duration and less.

Lawver, B. R.↗
NASA NTRS1985

Ignition characterization of LOX/hydrocarbon propellants

The results of an evaluation of the ignition characteristics of the gaseous oxygen (Gox)/Ethanol propellant combination are summarized. Ignition characterization was accomplished through the analysis, fabrication and testing of a spark initiated torch igniter and prototype 620 1bF thruster/igniter assembly. The igniter was tested over a chamber pressure range of 3.3 to 262 psia and a mixture ratio range of 0.4 to 40. The prototype thruster/igniter assembly was tested over the chamber pressure range of 74 to 197 psia and mixture ratio range of 0.778 to 3.29. Cold (-92 to -165 F) and ambient (44 to 80 F) propellant temperatures were used. Spark igniter ignition limits and thruster steady state and pulse mode, performance, cooling and stability data are presented. Spark igniter ignition limits are presented in terms of cold flow pressure, ignition chamber diameter and mixture ratio. Thruster performance is presented in terms of vacuum specific impulse versus engine mixture ratio. Gox/Ethanol propellants were shown to be ignitable over a wide range of mixture ratios. Cold propellants were shown to have a minor effect on igniter ignition limits. Thruster pulse mode capability was demonstrated with multiple pulses of 0.08 sec duration and less.

Lawver, B. R.↗
NASA NTRS1984

Ignition characterization of the GOX/ethanol propellant combination

This paper describes the results of a study to define the ignition characteristics and thruster pulse mode capabilities of the GOX/ethanol propellant combination. Ignition limits were defined in terms of mixture ratio and cold flow pressure using a spark initiated torch igniter. Igniter tests were run over a wide range of cold flow pressure, propellant temperature and mixture ratio. The product of cold flow pressure and igniter chamber diameter was used to correlate mixture ratio regimes of ignition and nonignition. Engine ignition reliability and pulse mode capability were demonstrated using a 620 lbF thruster with an integrated torch igniter. The nominal chamber pressure and mixture ratio were 150 psia and 1.8, respectively, thruster tests were run over a wide range of chamber pressures and mixture ratios. The feasibility of thruster pulse mode operation with the non-hypergolic GOX/ethanol propellant combination was demonstrated.

Lawver, B. R.↗
NASA NTRS1982

Test verification of LOX/RP-1 high-pressure fuel/oxidizer-rich preburner designs

Two fuel-rich and two oxidizer-rich preburner injectors are tested with LOX/RP-1 in an investigation of performance, stability and gas temperature uniformity over a chamber pressure range from 1292 to 2540 psia. Fuel-rich mixture ratios range from 0.238 to 0.367 and oxidizer-rich mixture ratios range from 27 to 48, and carbon deposition data are collected by measuring the pressure drop across a turbine simulator flow device. The oxidizer-rich testing demonstrates the feasibility of oxidizer-rich preburners, indicating equilibrium combustion as predicted, and the measured fuel-rich gas composition and C-asterisk performance are in excellent agreement with kinetic model predictions indicating kinetically-limited combustion.

Lawver, B. R.↗
NASA NTRS1982

Testing of fuel/oxidizer-rich, high-pressure preburners

Results of an evaluation of high pressure combustion of fuel rich and oxidizer rich LOX/RP-1 propellants using 4.0 inch diameter prototype preburner injectors and chambers are presented. Testing covered a pressure range from 8.9 to 17.5 MN/square meters (1292 to 2540 psia). Fuel rich mixture ratios ranged from 0.238 to 0.367; oxidizer rich mixture ratios ranged from 27.2 to 47.5. Performance, gas temperature uniformity, and stability data for two fuel rich and two ozidizer rich preburner injectors are presented for a conventional like-on-like (LOL) design and a platelet design injector. Kinetically limited combustion is shown by the excellent agreement of measured fuel rich gas composition and C performance data with kinetic model predictions. The oxidizer rich test results support previous equilibrium combustion predictions.

Lawver, B. R.↗
NASA NTRS1979

High performance N2O4/amine elements blowapart

An experimental and analytical program was conducted to develop an understanding of the mechanisms controlling hypergolic propellant reactive stream separation (RSS). RSS is a combustion induced phenomenon that results in striation of hypergolic propellant oxidizer and fuel sprays fans. This reduced intraelement mixing can influence thrust chamber performance, heat transfer, and stability. The program and product were the development of design criteria for coping with RSS to allow the design of high performance, stable injectors.

Lawver, B. R.↗
NASA NTRS1979

Photographic observation of reactive stream impingement

The theoretical and experimental study described was carried out to gain insight into the physical-chemical mechanism which controls hypergolic propellant reactive stream separation (RSS), with a view toward establishing design criteria. The approach employed was to observe photographically single element injector combustion of N2O4/MMH, N2O4/A-50, and N2O4/N2H4 propellants. Single element injectors were used to provide an unobstructed view of the impingement zone. Hot firings were conducted in a special photographic chamber. High-speed color motion pictures were used to identify the occurrence of RSS. Four conventional unlike doublet injectors, two conventional triplet injectors, and three platelet injectors were used in the tests. A simulation of the Space Shuttle platelet injector was included. The most important design criterion derived from the study states that the element should be designed to avoid transitions between mixed and separated modes within the engine operational envelope.

Lawver, B. R.↗
NASA NTRS1977

High performance N2O4/amine elements: Blowapart

The mechanisms controlling hypergolic propellant reactive stream separation (RRS) were studied and used to develop design criteria for injectors free from both steady state RSS and cyclic propellant stream separation. This was accomplished through the analysis of single element injectors using N204/MMH propellants; the injectors were representative of the space shuttle orbit maneuvering engine and space tug applications. A gas phase/surface reaction mechanism which controls RSS was identified. Injector design criteria were developed, which defined a critical chamber pressure for those operating conditions above which RSS occurs. It was found that the amount of interfacial surface area at impingement is controlled by injector hydraulics.

Lawver, B. R.↗
NASA NTRS1974

High performance N204/amine elements: Blowapart

The work is reported which was conducted to define the mechanisms governing blowapart of hypergolic propellant through the design, fabrication, test, and analysis of single element injectors. Data were developed that show the parameters exhibiting a controlling influence over blowapart are the chamber pressure, orifice diameter, and propellant temperature. Mixing, popping (cyclic blowapart), low pressure separation, and high pressure separation were identified as modes of reactive impingement.

Lawver, B. R.↗