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A sustained-arc ignition system was developed for internal combustion engines. It produces a very-long-duration ignition pulse with an energy in the order of 100 millijoules. The ignition pulse waveform can be controlled to predetermined actual ignition requirements. The design of the sustained-arc ignition system is presented in the report.
The pyrogen igniter has the function to furnish a controlled, high temperature, high pressure gas to ignite solid propellant surfaces in a rocket motor. Present pyrogens consist of numerous inert components. The Standard Molded Pyrogen Igniter (SMPI) consists of three basic parts, a cap with several integrally molded features, an ignition pellet retainer plate, and a tube with additional integrally molded features. A description is presented of an investigation which indicates that the SMPI concept is a viable approach to the design and manufacture of pyrogen igniters for solid propellant rocket motors. For some applications, combining the structural and thermal properties of molded composites can result in the manufacture of lighter assemblies at considerable cost reduction. It is demonstrated that high strength, thin walled tubes with high length to diameter ratios can be fabricated from reinforced plastic molding compound using the displacement compression process.
Conditions were determined in a premixing prevaporizing fuel preparation duct at which ignition occurred. An air blast type fuel injector with nineteen fuel injection points was used to provide a uniform spatial fuel air mixture. The range of inlet conditions where ignition occurred were: inlet air temperatures of 600 to 1000 K air pressures of 180 to 660 kPa, equivalence ratios (fuel air ratio divided by stoichiometric fuel air ratio) from 0.12 to 1.05, and velocities from 3.5 to 30 m/s. The duct was insulated and the diameter was 12 cm. Mixing lengths were varied from 16.5 to 47.6 and residence times ranged from 4.6 to 107 ms. The fuel was no. 2 diesel. Results show a strong effect of equivalence ratio, pressure and temperature on the conditions where ignition occurred. The data did not fit the most commonly used model of auto-ignition. A correlation of the conditions where ignition would occur which apply to this test apparatus over the conditions tested is (p/V) phi to the 1.3 power = 0.62 e to the 2804/T power where p is the pressure in kPa, V is the velocity in m/e, phi is the equivalence ratio, and T is the temperature in K. The data scatter was considerable, varying by a maximum value of 5 at a given temperature and equivalence ratio. There was wide spread in the autoignition data contained in the references.
Although the subject of spontaneous ignition of liquid fuels has received considerable attention in the past, the role of fuel evaporation in the overall spontaneous ignition process is still unclear. A main purpose of this research is to carry out measurements of ignition delay times, using fuels of current and anticipated future aeronautical interest, at test conditions that are representative of those encountered in modern gas turbine engines. Attention is focused on the fuel injection process, in particlar the measurement and control of man fuel drop size and fuel-air spatial distribution. The experiments are designed to provide accurate information on the role of fuel evaporation processes in determining the overall ignition delay time. The second objective is to examine in detail the theoretical aspects of spontaneous ignition in order to improve upon current knowledge and understanding of the basic processes involved, so that the results of the investigation can find general and widespead application.
The effect of oxygen pressure on the ignition temperature of cylindrical 302 SS specimens ignited by a focused CW CO2 laser beam in a cool, static, oxygen environment was investigated. The ignition temperature was determined quantitatively from the specimen temperature history obtained from a fast response, near infrared, two-color pyrometer, which recorded the temperature history of a spot approximately 0.5 mm in diameter and located at the center of the cylindrical 302 SS specimen's top surface. Ignition and combustion occurred in three stages. The first stage began with specimen heating and ended with ignition, the second stage corresponded to combustion, and the third stage represented cooling. The physical changes that occurred in the first stage are described.
An investigation of the pulse ignition characteristics of a 1 kW class arcjet using an inductive energy storage pulse generator with a pulse width modulated power converter identified several thruster and pulse generator parameters that influence breakdown voltage including pulse generator rate of voltage rise. This work was conducted with an arcjet tested on hydrogen-nitrogen gas mixtures to simulate fully decomposed hydrazine. Over all ranges of thruster and pulser parameters investigated, the mean breakdown voltages varied from 1.4 to 2.7 kV. Ignition tests at elevated thruster temperatures under certain conditions revealed occasional breakdowns to thruster voltages higher than the power converter output voltage. These post breakdown discharges sometimes failed to transition to the lower voltage arc discharge mode and the thruster would not ignite. Under the same conditions, a transition to the arc mode would occur for a subsequent pulse and the thruster would ignite. An automated 11 600 cycle starting and transition to steady state test demonstrated ignition on the first pulse and required application of a second pulse only two times to initiate breakdown.
An investigation of the pulse ignition characteristics of a 1 kW class arcjet using an inductive energy storage pulse generator with a pulse width modulated power converter identified several thruster and pulse generator parameters that influence breakdown voltage including pulse generator rate of voltage rise. This work was conducted with an arcjet tested on hydrogen-nitrogen gas mixtures to simulate fully decomposed hydrazine. Over all ranges of thruster and pulser parameters investigated, the mean breakdown voltages varied from 1.4 to 2.7 kV. Ignition tests at elevated thruster temperatures under certain conditions revealed occasional breakdowns to thruster voltages higher than the power converter output voltage. These post breakdown discharges sometimes failed to transition to the lower voltage arc discharge mode and the thruster would not ignite. Under the same conditions, a transition to the arc mode would occur for a subsequent pulse and the thruster would ignite. An automated 11 600 cycle starting and transition to steady state test demonstrated ignition on the first pulse and required application of a second pulse only two times to initiate breakdown.
Ignition of rich hydrogen-oxygen mixture in a monolithic catalytic reactor is studied using a transient combustion model. The model assumes a quasi-steady gas phase and a thermally-thin substrate. The ignition time lag is due to the thermal inertia of the substrate. One-step global chemical reaction is assumed both on the surface and in the gas phase. An effectiveness factor is introduced to account for the transition from a kinetically-limited catalytic surface reaction to a diffusion-limit one during ignition. Results presented include a catalytic ignition boundary, ignition delay time and the transient response in the catalytic bed.
Testing to evaluate Redesigned Solid Rocket Motor igniter heater electromagnetic interference (EMI) effects on the Safe and Arm (S and A) device was completed. It was suspected that EMI generated by the igniter heater and it's associated electromechanical relay could cause a premature firing of the NASA Standard Initiators (NSIs) inside the S and A. The maximum voltage induced into the NSI fire lines was 1/4 of the NASA specified no-fire limit of one volt (SKB 26100066). As a result, the igniter heaters are not expected to have any adverse EMI effects on the NSIs. The results did show, however, that power switching causes occasional high transients within the igniter heater power cable. These transients could affect the sensitive equipment inside the forward skirt. It is therefore recommended that the electromechanical igniter heater relays be replaced with zero crossing solid state relays. If the solid state relays are installed, it is also recommended that they be tested for EMI transient effects.
The ignition of carbon steel and 316 and 304 stainless steels caused by the impact of low-velocity particles (a standard mixture consisting of 2 g of iron and 3 g of inert materials) in gaseous oxygen was investigated using NASA/White Sands Test Facility for the ignition test, and a subsonic particle impact chamber to accelerate the particles that were injected into flowing oxygen upstream of the target specimen. It was found that the oxygen velocities required to ignite the three alloys were the same as that required to ignite the particle mixture. Ignition occurred at oxygen velocities greater than 45 m/sec at 20 to 24 MPa and was found to be independent of pressure between 2 and 30 MPa. Comparison of the present results and the past results from Wegener (1964) with the Compressed Gas Association (CGA) oxygen velocity limits for safe operations indicates that the CGA limits may be excessively conservative at high pressures and too liberal at low pressures.
Space Transportation System (STS-26) was launched from KSC pad 39B. Two of the Redesigned Solid Rocket Motors (RSRM) were part of the launch system and are designated RSRM-1A and RSRM-1B. Each of these motors were initiated by an 1U75164 igniter. The post flight condition of the igniters and associated components is documented. The overall performance of the igniter components was excellent. No damage or heat affected areas were noted. The sealing elements of the igniter functioned as expected with no evidence of erosion or blowby. The thermal protection system protected all areas adequately. No excessive erosion was noted. Corrosion was found in the special bolt holes in the igniter chamber. The corrosion will not affect refurbishment of the chamber. Beginning with flight 5 grease has been added to the chamber holes to prevent this erosion.
A theory of evaporation, ignition, and burning of moderately dense spherical drop clusters has been developed. The theory takes into account burning of premixed air and fuel internal to the cluster at ignition and subsequent burning of fuel emitted from the cluster by a flame sheet surrounding it. The model considers interdrop interaction, momentum exchange between drops and gas, and turbulent exchange processes between the cluster and its surroundings. Calculations are performed for varying initial air-to-fuel-mass ratios, initial cluster radii, ambient gas temperatures and initial drop temperatures. Results are presented for ratios of fuel mass burned to fuel mass lost from the cluster between drop ignition and drop disappearance, fuel burned fractions at ignition and at the moment of drop disappearance, and jump conditions at ignition.
Results from an experimental study of laser induced spark ignition of methane-oxygen mixtures are presented. The experiments were conducted at atmospheric pressure and 296 K under laminar pre-mixed and turbulent-incompletely mixed conditions. A pulsed, frequency doubled Nd:YAG laser was used as the ignition source. Laser sparks with energies of 10 mJ and 40 mJ were used, as well as a conventional electrode spark with an effective energy of 6 mJ. Measurements were made of the flame kernel radius as a function of time using pulsed laser shadowgraphy. The initial size of the spark ignited flame kernel was found to correlate reasonably well with breakdown energy as predicted by the Taylor spherical blast wave model. The subsequent growth rate of the flame kernel was found to increase with time from a value less than to a value greater than the adiabatic, unstretched laminar growth rate. This behavior was attributed to the combined effects of flame stretch and an apparent wrinkling of the flame surface due to the extremely rapid acceleration of the flame. The very large laminar flame speed of methane-oxygen mixtures appears to be the dominant factor affecting the growth rate of spark ignited flame kernels, with the mode of ignition having a small effect. The effect of incomplete fuel-oxidizer mixing was found to have a significant effect on the growth rate, one which was greater than could simply be accounted for by the effect of local variations in the equivalence ratio on the local flame speed.
The use of laser-induced spark ignition in liquid-propellant rocket engines is investigated by studying the ignition of mixtures of methane and oxygen. Specific attention is given to minimum ignition energy, spark efficiency, the effect of ignition energy on flame kernels, the effect of laser wavelength, and comparisons with electrode-spark ignition. A pulsed frequency-doubled Nd:YAG laser is tested at atmospheric pressure and 296 K under both laminar-premixed and turbulent-incompletely-mixed conditions. Laser sparks of 10 and 40 mJ and an electrode spark of 6 mJ are measured for flame-kernel radius as a function of time with pulsed laser shadowgraphy. Initially, the flame-kernel size is similar to those predicted by the Taylor spherical-blast-wave model, and subsequent growth is characterized by rapid acceleration. The growth rate is significantly affected by the effect of incomplete fuel-oxidizer mixing.
Experimental and analytical investigations focusing on secondary atomization and ignition characteristics of aluminum/liquid hydrocarbon slurry propellants were conducted. Experimental efforts included the application of a laser-based, two-color, forward-scatter technique to simultaneously measure free-flying slurry droplet diameters and velocities for droplet diameters in the range of 10-200 microns. A multi-diffusion flame burner was used to create a high-temperature environment into which a dilute stream of slurry droplets could be introduced. Narrowband measurements of radiant emission were used to determine if ignition of the aluminum in the slurry droplet had occurred. Models of slurry droplet shell formation were applied to aluminum/liquid hydrocarbon propellants and used to ascertain the effects of solids loading and ultimate particle size on the minimum droplet diameter that will permit secondary atomization. For a 60 weight-percent Al slurry, the limiting critical diameter was predicted to be 34.7 microns which is somewhat greater than the 20-25 micron limiting diameters determined in the experiments. A previously developed model of aluminum ignition in a slurry droplet was applied to the present experiments and found to predict ignition times in reasonable agreement with experimental measurements. A model was also developed that predicts the mechanical stress in the droplet shell and a parametric study was conducted. A one-dimensional model of a slurry-fueled rocket combustion chamber was developed. This model includes the processes of liquid hydrocarbon burnout, secondary atomization, aluminum ignition, and aluminum combustion. Also included is a model for radiant heat transfer from the hot aluminum oxide particles to the chamber walls. Exercising this model shows that only a modest amount of secondary atomization is required to reduce residence times for aluminum burnout, and thereby maintain relatively short chamber lengths. The model also predicts radiant heat transfer losses to the walls to be only approximately 3 percent of the fuel energy supplied. Additional work is required to determine the effects of secondary atomization on two-phase losses in the nozzle.
A laser ignition system is proposed for the Combustion Experiment Module on an orbiting spacecraft. The results of a design study are given using the scheduled 'Flame Ball Experiment' as the design guidelines. Three laser ignition mechanisms and wavelengths are evaluated. A prototype laser is chosen and its specifications are given, followed by consideration of the beam optical arrangement, the ignition power requirement, the laser ignition system weight, size, reliability, and laser cooling and power consumption. Electromagnetic interference to the onboard electronics caused by the laser ignition process is discussed. Finally, ground tests are suggested.
The ignition and combustion of Al, Mg, and Al/Mg alloy particles in 99 percent O2/1 percent N2 mixtures is investigated at high temperatures and pressures for rocket engine applications. The 20-micron particles contain 0, 5, 10, 20, 40, 60, 80, and 100 wt pct Mg alloyed with Al, and are ignited in oxygen using the reflected shock in a single-pulse shock tube near the endwall. Using this technique, the ignition delay and combustion times of the particles are measured at temperatures up to 3250 K as a function of Mg content for oxygen pressures of 8.5, 17, and 34 atm. An ignition model is developed that employs a simple lumped capacitance energy equation and temperature and pressure dependent particle and gas properties. Good agreement is achieved between the measured and predicted trends in the ignition delay times.
Motion pictures were taken at 1,850 frames per second of the spray penetration and combustion occurring in the N.A.C.A. combustion apparatus arranged to operate as a compression-ignition engine. Indicator cards were taken simultaneously with the motion pictures by means of the N.A.C.A. optical indicator. The motion pictures showed that when ignition occurred during injection it started in the spray envelope. If ignition occurred after injection cut-off, however, and after considerable mixing had taken place, it was impossible to predict where the ignition would start. The pictures also showed that ignition usually started at several points in the combustion chamber. With this apparatus, as the injection advance angle increased from 0 degrees to 40 degrees before top center, the rate of flame spread increased and the duration of the burning decreased.