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Friction and compression induced ignition prevented by heat-dissipation layer. Thin layer of copper placed inside surfaces of pipes and other parts in contact with flowing liquid oxygen, fluorine, or other oxidizers. Copper layer of only 0.001 to 0.005 inch (0.0254 to 0.127 mm) ensures heat away from hotspots thereby discouraging ignition. Copper layer increases oxygen ignition threshold of pipe by 5 to 20 times.
Topics discussed are related to elementary reactions, reaction mechanisms and modeling, laminar flames, flame chemistry, turbulent reacting shear flows, turbulent premixed flames, turbulent combustion measurements, continuous combustors, detonation, detonation and explosion, heterogeneous detonation, propellant combustion, fire-ignition and thermal degradation, fire-flame spread and burning, fire-modeling, spray combustion, and droplet combustion. Coal combustion kinetics and mechanisms are considered along with coal combustion mechanisms and pyrolysis, coal combustion techniques, NOx in coal combustion, gaseous pollutants, soot and PAH, soot and inorganic pollutants, I.C. engine combustion, and ignition and extinction. Attention is given to intricate paths and simple steps in chemical kinetics and combustion, the formation of polycyclic aromatic hydrocarbons by combustion, turbulent flame structure and speed in spark ignition engines, and unresolved problems in SOx, NOx, and soot control in combustion.
A low gravity exothermic heating/cooling apparatus is disclosed for processing materials in space which includes an insulated casing and a sample support carried within the casing which support a sample container. An exothermic heat source includes a plurality of segments of exothermic material stacked one upon another to produce a desired temperature profile when ignited. The sample container is arranged within the core of the stacked exothermic heating material. Igniters are spaced vertically along the axis of the heating material to ignite the exothermic material at spaced points to provide total rapid burn and release of heat. To rapidly cool and quench the heat, a source of liquid carbon dixoide is provided which is conveyed through a conduit and a metering orifice into a distribution manifold where the carbon dioxide is gasified and dispersed around the exothermic heating material and the sample container via tubes for rapidly cooling the material sample.
The present conference on combustion phenomena considers topics in automotive engine combustion, turbulent reacting flows, the modeling of practical combustion systems, reaction kinetics, combustion-generated particulates, combustion diagnostics, coal combustion process characteristics, fire-related phenomena, explosion/detonation phenomena, spray combustion, ignition/extinction, laminar flames, pollutant formation processes, practical combustor devices, and rocket propellant combustion. Attention is given to the contributions of combustion science to piston engine design, modeling and measurement techniques for turbulent combustion, the specific effects of energy, collisions, and transport processes in combustion chemistry kinetics, the formation of large molecules, particulates and ions in premixed hydrocarbon flames, the application of laser diagnostics to combustion systems, spark ignition energies for dust-air mixtures, the controlling mechanisms of flow-assisted flame spread, the ignition and combustion of coal-water slurries, spontaneous ignition of methane, turbulent and accelerating dust flames, and the temperature sensitivity of double base propellants.
A series of direct connect combustor tests was conducted to determine the effect of silane concentration on the supersonic combustion characteristics of silane/methane mixtures. Shock tube ignition delay data indicated more than an order of magnitude reduction in ignition delay times for both 10 and 20 percent silane/methane mixtures as compared to methane. The ignition delay time of the 10 percent mixture was only a factor of 2.3 greater than that of the 20 percent mixture. Supersonic combustion tests were conducted with the fuel injected into a model scramjet combustor. The combustor was mounted at the exit of a Mach 2 nozzle and a hydrogen fired heater was used to provide a variation in test gas total temperature. Tests using the 20 percent silane/methane mixture indicated considerable combustion enhancement when compared to methane alone. This mixture had an autoignition total temperature of 1650 R. This autoignition temperature can be contrasted with 2330 R for hydrogen and 1350 R for a 20 percent silane/hydrogen mixture in similar hardware. Methane without the silane additive did not autoignite in this configuration at total temperatures as high as 3900 R, the maximum temperature at which tests were conducted. Supersonic combustion tests with the silane concentration reduced to 10 percent indicated little improvement in combustion performance over pure methane. The addition of 20 percent silane to methane resulted in a pyrophoric fuel with good supersonic combustion performance. Reducing the silane concentration below this level, however, yielded a less pyrophoric fuel that exhibited poor supersonic combustion performance.
Nasa Lewis' studies of complex reacting systems at high temperature are discussed. The changes which occur are the result of many different chemical reactions occurring at the same time. Both an experimental and a theoretical approach are needed to fully understand what happens in these systems. The latter approach is discussed. The differential equations which describe the chemical and thermodynamic changes are given. Their solution by numerical techniques using a detailed chemical mechanism is described. Several different comparisons of computed results with experimental measurements are also given. These include the computation of (1) species concentration profiles in batch and flow reactions, (2) rocket performance in nozzle expansions, and (3) pressure versus time profiles in hydrocarbon ignition processes. The examples illustrate the use of detailed kinetic computations to elucidate a chemical mechanism and to compute practical quantities such as rocket performance, ignition delay times, and ignition lengths in flow processes.
The density at which electron capture can ignite oxygen burning in 8-10 solar mass stars is investigated within the uncertainties of the semiconvection theory. The Ledoux criterion for convection is adopted in order to take into account the maximum effect of heating due to electron capture. It is found that oxygen burning is not ignited by electron captures on Mg-24 and Na-24 at the central density of about 4 x 10 to the 9th g/cu cm, but by capture on Ne-20 at about 1 x 10 to the 10th g/cu cm. This ignition density is 2.5 times lower than in the case with the Schwarzschild criterion, but is still high enough to induce a collapse by electron capture rather than a total explosion of the core. When the oxygen-burning front arrives at M(r) of roughly 0.3 solar mass, the contraction reaches a quasi-dynamic collapse.
It is shown here that the behavior of nuclear burning on accreting neutron stars depends on the mass accretion rate per unit area sigma-dot, the temperature T, and the column density of accreted matter sigma, and is thus intrinsically a three-dimensional problem. The concepts of a nuclear fuel surface and an ignition surface in (sigma-dot, T, sigma) space is introduced. It is shown that certain cuts taken through the ignition surface correspond to particular physical situations. It is found that at high temperatures a succession of hydrogen flashes leads eventually to a pure helium flash. At low temperatures, compressional heating completely determines the burst behavior. Burst sources are expected to show highly erratic behavior if they lie near the boundaries of the H and He ignition surfaces, resolving a major discrepancy between the thermonuclear flash model and observations.
An analysis is presented of laboratory experiments that have shown that hollow plasma sources can provide low-impedance contact with a background plasma. In particular, some experiments show an report an 'ignited mode' of electron collection using plasma sources where almost an ampere was collected at a potential of 100 volts. The experimental results are compared with theoretical and computational models of plasma layers. Preliminary results indicate that the measured potential profiles and collected currents during the 'ignited mode' correspond to an increase in the background plasma thermal current of more than an order of magnitude. The results imply that the 'ignited mode' results are substantially impacted by the laboratory electron source and the tank size. Only experiments performed in space will be able to provide the correct boundary conditions for reliably testing high electron collection by plasma contactors.
Qualification testing of the Redesigned Solid Rocket Motor improved igniter-to-case joint and field joint heaters was conducted on the fired TEM-04 static test motor and was completed on 7 Sep. 1989. The purpose of the test was to certify the installation and performance of the improved joint heaters for use on flight motors. The changes incorporated in the improved heaters improve durability and should reduce handling damage. The igniter-to-case joint and field joint primary heater circuits were subjected to five 20-hr ON cycles. The heater redundant circuits were then subjected to one 20-hr ON cycle. Voltage, current, and temperature set point values were maintained within the specified limits for both heaters during each ON cycle. When testing was complete, both heaters were removed and inspected. No discolorations or any other anomalies were found on either of the heaters. Based on the successful completion of this test, it is recommended that the improved igniter-to-case joint and field joint heaters be used on future flight motors.
The Redesigned Solid Rocket Motors (RSRM) of the Space Transportation System have three field joints that are protected by the Joint Protection Systems (JPS). The igniter heater was mounted on the igniter flange. This report documents the performance of the JPS and igniter heaters on the pad and the post-flight condition of the JPS components. All observations that were written up as Squawks and/or Problem Reports are also discussed. The primary heaters performed satisfactorily and maintained the field joint temperatures within the required temperature range. A secondary heater failed Dielectric Withstanding Voltage (DWV) test during the joint closeout prior to launch. This heater was not used, however, since the primary heater functioned properly. Post-test inspection revealed that pin A of the heater power cable was shorted to the connector shell. Design changes have been implemented to resolve the heater power cable problem. All field joint assemblies met all of the performance requirements.
The performance is assessed of the 360H005, Fifth flight, Redesigned Solid Rocket Motors (RSMR) in respect to joint sealing issues as seen from post flight inspection of the seals and sealing surfaces. The factory joint disassembly inspections have resumed for 360H005. The new factory joint grease application is in effect and now can be assessed during the disassembly process. The RSRM is illustrated consisting of capture feature field joints as is the J-joint insulation configuration. The nozzle-to-case joint design is also illustrated, which includes 100, 7/8 inch radial bolts in conjunction with a wiper O-ring and modified insulation design. The ignition system seals and a cross section of the igniter are illustrated. The configuration of all the internal nozzle joints are also shown. The postflight inspection of both motors showed the seal components to be in excellent condition except for the indentation found on the inner primary seal of the right hand inner igniter gasket, aft face. Detailed inspection results, and inspections performed by the O-ring Inspection Team are presented.
Flight motor set 360L008 was launched as part of NASA space shuttle mission STS-32R. As with all previous redesigned solid rocket motor launches, overall motor performance was excellent. All ballistic contract end item specification parameters were verified with the exception of ignition interval and rise rates, which could not be verified due to elimination of developmental flight instrumentation. But the available low sample rate data showed nominal propulsion performance. All ballistic and mass property parameters closely matched the predicted values and were well within the required contract end item specification levels that could be assessed. All field joint heaters and igniter joint heaters performed without anomalies. Redesigned field joint heaters and the redesigned left-hand igniter heater were used on this flight. The changes to the heaters were primarily to improve durability and reducing handling damage. Evaluation of the ground environment instrumentation measurements again verified thermal mode analysis data and showed agreement with predicted environmental effects. No launch commit criteria violation occurred. Postflight inspection again verified superior performance of the insulation, phenolics, metal parts, and seals. Postflight evaluation indicated both nozzles performed as expected during flight. All combustion gas was contained by insulation in the field and case-to-nozzle joints. Recommendations were made concerning improved thermal modeling and measurements. The rationale for these recommendations and complete result details are presented.
Test apparatus and method developed to measure ignition capability of percussion primers. Closely simulates actual conditions and interfaces encountered in such applications as in munitions and rocket motors. Ignitability-testing apparatus is small bomb instrumented with pressure transducers. Sizes, shapes, and positions of bomb components and materials under test selected to obtain quantitative data on ignition.
The performance of the field joint protection system (FJPS) of the Space Shuttle Production Verification Motor 1 (PV-1), as evaluated by postfire hardware inspection. Compliance with the specifications is shown for the FJPS assembly and components. The simplified FJPS and field joint heaters performed nominally, maintaining all joint seal temperatures within the required range. One anomally was noted on the igniter-to-case joint heater during postfire inspection. The heater buckled off the surface in two areas, resulting in two hot spots on the heater and darkened heater insulation. The condition did not affect heater performance during ignition countdown and all igniter seals were maintained within required temperature limits.
Flight motor set 360T010 was launched at approximately 7:34 a.m. CST (090:114:12:33:50.990 GMT) on 24 Apr. 1990 after one launch attempt (attempt on 10 Apr. 1990 was scrubbed following an indication of erratic operation of the Orbiter No. 1 Auciliary Power Unit No. 1). There were no problems with the solid rocket motor launches, overall motor performance was excellent. There were no debris concerns from either motor. Nearly all ballistic contract end item specification parameters were verified with the exception of ignition interval, pressure rise rate, and ignition time thrust imbalance. These could not be verified due to elimination of developmental flight instrumentation on 360L004 (STS-30R) and subsequent, but low sample rate data that were available showed nominal propulsion performance. All ballistic and mass property parameters that could be assessed closely matched the predicted values and were well within the required contract end item specification levels. All field joint heaters and igniter joint heaters performed without anomalies. Evaluation of the ground environment instrumentation measurements again verified thermal model analysis data and showed agreement with predicted environmental effects. No launch commit criteria violations occurred. Postflight inspection again verified nominal performance of the insulation, phenolics, metal parts, and seals. Postflight evaluation indicated both nozzles performed as expected during flight. All combustion gas was contained by insulation in the field and case-to-nozzle joints.
An extensive program to improve the operation of the Ames 16 Inch Shock Tunnel was carried out. This included reduction of driver slosh wave amplitudes and detonation risk by the use of premixed He/O2 gas, longer wait times between the last gas load and driver gas ignition, an improved gas loading sequence, the use of four instead of one ignition wire, and the use of lower ignition wire voltages. Successful operation of the tunnel at pressures of 2000-6000 psi and enthalpies up to 12,000 J/gm was achieved. A new diaphragm breaking technique, self break on the driver combustion pressure rise, was tested and found to produce clean breaks over the full pressure range of the tunnel. Improvements were made to the driver gas loading manifold and a preliminary design was made of a gas mixing system which mixes all three gases on the fly just before injection into the driver. Other improvements as well as tests are discussed.
Fire safety is an important issue in the current space program; ignition in low-g needs to be studied. The reduction in the gravitational acceleration causes changes in the ignition process. This paper examines the effect of gravity on natural convection, which is one of the important parameters in the ignition process. The NASA-Lewis 2.2 Second Drop Tower provided the low-gravity environment for the experiments. A series of experiments was conducted to measure the temperature of a small copper plate which was heated by a high intensity lamp. These experiments verified that in low-gravity the plate temperature increased faster than in the corresponding 1-g cases, and that the natural convection coefficient rapidly decreased in the low-gravity environment.