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Jet pluming effect on static stability of five cone cylinder-flare finned rocket models at Mach numbers of 4, 5, and 6 and static pressure ratios up to 26,000
Martian soil and surface changes under rocket engine exhaust loads
Calculation of crew radiation dosage from uranium gas nuclear rocket engine exhaust backflow of fission products
Spectral scanning method for determining temperature profile of jet- or rocket-engine exhaust stream by gas radiation and transmittance measurements, discussing radiometric errors effects
The spectral-scanning method may be used to determine the temperature profile of a jet- or rocket-engine exhaust stream by measurements of gas radiation and transmittance, at two or more wavelengths. A single, fixed line of sight is used, using immobile radiators outside of the gas stream, and there is no interference with the flow. At least two sets of measurements are made, each set consisting of the conventional three radiometric measurements of absorption-emission pyrometry, but each set is taken over a different spectral interval that gives different weight to the radiation from a different portion of the optical path. Thereby, discrimination is obtained with respect to location along the path. A given radiometric error causes an error in computed temperatures. The ratio between temperature error and radiometric error depends on profile shape, path length, temperature level, and strength of line absorption, and the absorption coefficient and its temperature dependency. These influence the choice of wavelengths, for any given gas. Conditions for minimum temperature error are derived. Numerical results are presented for a two-wavelength measurement on a family of profiles that may be expected in a practical case of hydrogen-oxygen combustion. Under favorable conditions, the fractional error in temperature approximates the fractional error in radiant-flux measurement.
A cooled-gas pyrometer designed for application in a hypersonic research engine program was fabricated and tested. Design and operational considerations and calibration data are presented. The probe was tested in a rocket-engine exhaust stream operating at Mach 2 and 2300 K. Test temperature measurements agreed to within 2 percent with a radiation shielded thermocouple probe.
The development of an acid base gas-phase reaction system which utilizes anhydrous ammonia as the reactant to remove nitrogen dioxide from hydrazine-nitrogen tetroxide rocket combustion exhaust is reported. This reaction reduced NO2 levels in exhaust emissions so that the resulting stack emission is completely white instead of the earlier observed typical reddish-brown coloration. Preliminary analyses indicate the importance of reaction time and ammonia concentration on removal efficiency and elimination of the health hazard to individuals with respiratory problems.
A gamma ray spectrometer is described for use in intense radiation fields such as those in the vicinity of a rocket engine exhaust. A collimated radiation beam is Compton scattered toward shielded spectrometers to reduce the energy and intensity of the radiation and is energy selective among the spectrometers. The scattering targets are changeable to control the percentage of the radiation scattered. Sum-Compton coincidence techniques are employed for data selection.
The spectral-scanning method may be used to determine the temperature profile of a jet- or rocket-engine exhaust stream by measurements of gas radiation and transmittance at two or more wavelengths. A single, fixed line of sight is used, using immobile radiators and radiometers outside of the gas stream, and there is no interference with the flow. A given radiometric error causes an error in computed temperatures. The ratio between temperature error and radiometric error depends on profile shape, path length, temperature level, the strength of line absorption, and the absorption coefficient and its temperature coefficient. These influence the choice of wavelengths for any given gas. Conditions for minimum temperature error are derived. Numerical results are presented for a two wavelength measurement on a family of profiles that may be expected in a practical case of H2 - O2 combustion. Under favorable conditions, the fractional error in temperature approximates the fractional error in radiant-flux measurement.
The screech phenomenon observed in a one-sixtieth scale model space shuttle test of the solid rocket booster exhaust flow noise has been investigated. A critical review is given of the cold flow test data representative of Space Shuttle launch configurations to define those parameters which contribute to screech generation. An acoustic feedback mechanism is found to be responsible for the generation of screech. A simple equation which permits prediction of screech frequency in terms of basic testing parameters such as the jet exhaust Mach number and the separating distance from nozzle exit to the surface of model launch pad is presented and is found in good agreement with the test data. Finally, techniques are recommended to eliminate or reduce the screech.
The contamination and degradation of externally exposed sensitive experimental and functional surfaces of the Skylab by the self-induced atmosphere surrounding the vehicle are studied. It is shown that the major contamination sources of long-term deposits on Skylab are primarily outgassing from nonmetallic coatings and rocket engine exhaust products. Analysis showed that the characteristic decay time of the outgassing source rate may be as long as 4100 hr, and that essentially 20% of RCS deposits that stick at temperatures of 0 C will remain after sublimating at a rate whose characteristic decay time is 72 hr.
A computer program to calculate the temperature profile of a flame or hot gas was presented in detail. Emphasis was on profiles found in jet engine or rocket engine exhaust streams containing H2O or CO2 radiating gases. The temperature profile was assumed axisymmetric with an assumed functional form controlled by two variable parameters. The parameters were calculated using measurements of gas radiation at two wavelengths in the infrared. The program also gave some information on the pressure profile. A method of selection of wavelengths was given that is likely to lead to an accurate determination of the parameters. The program is written in FORTRAN IV language and runs in less than 60 seconds on a Univac 1100 computer.
A multistage cascade impactor instrument employing piezoelectric crystals for real-time aerosol size distribution measurements is described in this paper. The instrument measures the aerosol mass concentration as a function of particle size in 10 size intervals ranging from 0.05 micrometers to 25 micrometers in diameter. The particles are then analyzed for elemental composition and morphology by means of scanning electron microscopy (SEM). Examples of particle size distributions in rocket motor exhaust plumes measured with the instrument are presented together with the elemental composition and morphology obtained by SEM.
Sliding-gate valve originally developed for rocket exhaust-gas ducts is sealed by a Belleville spring. It is simple, compact, and operates over a wider range of temperatures than conventional O-ring sealed valves.
Changes in the atmospheric composition, particularly through the condensation of rocket vehicle exhaust, caused by the flights of 400 heavy lift launch vehicles (HLLV) to carry crews and materials into space to build a satellite solar power system (SPS) were examined. Attention was given to the formation of mesospheric contrails and clouds. A one-dimensional model was used to formulate the photochemistry and vertical transport of water vapor, its nucleation into an ice cloud, and the microphysical development of the cloud. Considering one HLLV launch per day for a decade, it is projected that the upper atmosphere water vapor concentration would be increased by 10-20%, thereby augmenting the size and opacity of natural noctilucent clouds by 50%. No climatological consequences are foreseen from the clouds, although spectacular noctiluminescent cloud displays are thought to be possible.
A solution was generated for the physical properties of the Shuttle RCS 4000 N (900 lb) rocket engine exhaust plume flowfield. The modeled exhaust gas consists of the five most abundant molecular species, H2, N2, H2O, CO, and CO2. The solution is for a bare RCS engine firing into a vacuum; the only additional hardware surface in the flowfield is a cylinder (=engine mount) which coincides with the nozzle lip outer corner at X = 0, extends to the flowfield outer boundary at X = -137 m and is coaxial with the negative symmetry axis. Continuum gas dynamic methods and the Direct Simulation Monte Carlo (DSMC) method were combined in an iterative procedure to produce a selfconsistent solution. Continuum methods were used in the RCS nozzle and in the plume as far as the P = 0.03 breakdown contour; the DSMC method was used downstream of this continuum flow boundary. The DSMC flowfield extends beyond 100 m from the nozzle exit and thus the solution includes the farfield flow properties, but substantial information is developed on lip flow dynamics and thus results are also presented for the flow properties in the vicinity of the nozzle lip.
The potential benefits of simultaneously using hydrogen and oxygen as rocket engine coolants are described. A plug-and-spool rocket engine was examined at heat fluxes ranging from 9290 to 163,500 kW/sq m, using a combined 3-D conduction/advection analysis. Both counter flow and parallel flow cooling arrangements were analyzed. The results indicate that a significant amount of heat transfer to the oxygen occurs, reducing both the hot side wall temperature of the rocket engine and also reducing the exit temperature of the hydrogen coolant. In all heat flux and coolant flow rates examined, the total amount of heat transferred to the oxygen was found to be largely independent of the oxygen coolant flow direction. At low heat flux/low coolant flow (throttled) conditions, the oxygen coolant absorbed more than 30 percent of the overall heat transfer from the rocket engine exhaust gasses. Also, hot side wall temperatures were judged to decrease by approximately 120 K in the throat area and up to a 170 K combustion chamber wall temperature reduction is expected if dual cooling is applied. The reduction in combustion chamber wall temperatures at throttled conditions is especially desirable since tha analysis indicates that a double temperature maxima, one at the throat and another in the combustion chamber, occurs with a traditional hydrogen cooled only engine. Conversely, a dual cooled engine essentially eliminates any concern for overheating in the combustion chamber.
The Diagnostics Testbed Facility (DTF) at NASA's John C. Stennis Space Center in Mississippi was designed to provide a testbed for the development of rocket engine exhaust plume diagnostics instrumentation. A 1200-lb thrust liquid oxygen/gaseous hydrogen thruster is used as the plume source for experimentation and instrument development. Theoretical comparative studies have been performed with aerothermodynamic codes to ensure that the DTF thruster (DTFT) has been optimized to produce a plume with pressure and temperature conditions as much like the plume of the Space Shuttle Main Engine as possible. Operation of the DTFT is controlled by an icon-driven software program using a series of soft switches. Data acquisition is performed using the same software program. A number of plume diagnostics experiments have utilized the unique capabilities of the DTF.