Showing papers on "Burn rate (chemistry) published in 1990"

Analysis of the Flight Performance of the 155 mm M864 Base Burn Projectile

Abstract: : An engineering model has been developed to compute the flight performance of the M864 base burn projectile. This model includes the coupled performance of the gas generator, effects of injected mass flow on the aerodynamics, and a modified point mass trajectory simulation. The gas generator model is based on measured burn rates and basic fluid dynamics. The discharge rate of the generator is calibrated against laboratory experiments. Effects of spin on burn rate are deduced from comparison of analysis with spin fixture tests. Linear and first order nonlinear effects of mass injection on base pressure are the basis for evaluation of base drag. Navier-Stokes solutions near the base with air injection provides essential data. Correlation equations predict base pressure as a function of Mach number, injection rate, and propellant gas temperature. Temperature effects on base pressure are a unique feature of the analysis. Keywords: Injection of combustion products, Base burn, Base bleed, Spin, Trajectory, Navier-Stokes computations, Solid propellant.

Driving of axial acoustic fields by sidewall stabilized diffusion flames.

Abstract: This paper describes recent results obtained in a study concerned with the elucidation of the mechanisms which drive instabilities in solid propellant rocket motors. In this study, the interaction between diffusion flames, stabilized on the side wall of a long rectangular duct, and an axial acoustic field was investigated experimentally. The behavior of the flame under a variety of test conditions was investigated using high speed shadowgraph movies, a light intensified imaging system, and C-H flame radiation measurements. The high speed cinematography and intensified imaging system showed that the excitation of acoustic waves produced axial and transverse flame oscillations with the frequency of the imposed waves. The flame radiation measurements revealed that the presence of an acoustic field produced space dependent oscillatory reaction and heat release rates which depend upon the characteristics of the flame and the excited acoustic field. Measurements of the space dependence of the heat release rates ., showed that at a given instant some sections of the flame drive the acoustic field while other damp it. The net effect of the flame upon the acoustic field depends upon the relative magnitudes of these driving and damping regions, Considerations of the physics of the problem suggest that both the acoustic pressure and velocity oscillations contribute to the observed flame behavior. This paper describes recent results obtained in an ongoing, AFOSR sponsored, investigation of the mechanisms which drive axial instabilities in solid propellant rocket motors. Undesirable combustion instabilities occur when energy supplied by the combustion process to flow disturbances exceeds the energy extracted by loss mechanisms (e.9.. viscous dissipation), resulting in the excitation of large amplitude wave motions inside the combustor. Often, these combustor flow oscillations excite wave motions inside the * Graduate Research Assistant * * Research Engineer, Member AIAA t Reqents' Professor, Fellow AIAA 1 propellant grain, the motor casing and related systems. These wave motions generally produce undesirable side effects which may include mechanical failures of system components, modification of the propellant burn rate, vibrations in the control system, and so on. Since individually, or in combination, these effects can lead to mission failure, it is of utmost importance that capabilities for eliminating or reducing the onset of such instabilities be developed. This goal can be attained by developing means for reducing the driving provided by the combustion process and/or increasing the damping experienced by the waves, The study described herein has been concerned with developing of an understanding of the role that the gas phase portion of a solid propellant flame plays in the driving process. Solid propellant flames are extremely complex. They generally consist of a pyrolyzing solid propellant which supplies gaseous streams of fuel and oxidizer which burn in a complex myriad of premixed and diffusion flames1. A fraction of the heat released by these flames is fed back to the solid propellant to sustain the pyrolysis reactions. The complexity of these flames increases considerably during an instability when the various flame processes become unsteady, and the interaction of the flame with the combustor pressure oscillations produces periodic flame movements. These unsteady flame processes are generally accompanied by periodic heat release processes which supply the energy required for driving the instability. Thus, to provide an understanding of the processes which drive combustion instabilities, the characteristics of unsteady solid propellant flames must be understood. Ideally, one would want to tackle this problem by investigating the characteristics of an actual solid propellant burning under conditions which simulate those encountered in an unsteady rocket motor. However, to date, the extremely small dimensions of the gas and condensed phases2 which constitute a solid propellant flame (i.e., they are of the order of microns), the smoky nature of the flame, the high burn rate of the propellant (which limits the time available for conducting the experiment), and the limitations of existing measurement systems have prevented investigators from attaining this goal. Instead, investigators have resorted to the study of "idealized" flames, which possessed certain important features of actual solid propellant flames, under conditions which simulated those encountered in unstable solid propellant rocket motors. F o r example, Kumar et a13 used a porous plate burner to simulate the gas phase flame of a non-metallized composite propellant. Their study showed that the diffusive mixing of the oxidizer and fuel vapors controlled the extent of the gas phase combustion zone by affecting the heat transfer and hence the propellant burning rate. Brown et a14 studied ammonium perchlorate (AP) -binder combustion in steady and high acceleration environments by using an oxidizer-binder sandwich simulation. They found that the combustion process at low pressures is laminar and that the fuel is burned in a diffusion flame in the vicinity of the interface between the binder and the AP. At higher pressures, the combustion process appears to be turbulent and consist of premixed (i.e., AP deflagration) and diffusion flame regions. While these studies provide considerable insight into the characteristics of actual, steady, solid propellant flames, they have not considered the complex issues associated with the unsteady combustion of these propellants. These studies have indicated, however, that studies concerned with the driving of instabilities by actual solid propellant flames will have to investigate the contributions from both the diffusion and premixed flames which are present in the gas phase flames of solid propellant propellants. Recently, the contribution of the premixed gas phase flames of solid propellants to the driving of axial instabilities were investigated theoretically and experimentally by the authors and coworker.^^-^. Specifically, the characteristics of a premixed flat flame stabilized on the side wall of duct in which an axial acoustic field had been excited were investigated. Special emphasjs was Placed on elucidating the mechanisms through the investigated flame added energy to the acoustic field. These studies shoved that the interaction of the premixed flat flame with the acoustic field produced oscillatory reaction rate and movement of the flame relative to the side wall, It was also shown that driving provided by the investigated premixed flame is acoustically equivalent to that provided by a combination of a monopole and a dj.pole acoustic sources, The study described in this paper represents a continuation of the above described premixed flame Studies. It focuses on the determination of the characteristics and acoustic driving provi'ded by one o r more diffusion flames stabilized on the side wall of an acoustically excited rectangular duct, see Fig. 1. Specifically, it is concerned with the mechanisms through which diffusion flames drive axial acoustic waves, and the magnitude of this driving relative to the magnitude of the driving provided by the previously investigated premixed flames. While this study uses gaseous diffusion flames to investigate the driving by actual solid propellant flames, it should be pointed out that the investigated flame configurations possess important features similar to those found in solid propellant flames. For example, both flames interact with the thermal and velocity acoustic boundary layers which exist near burning solid propellant surfaces, and in both cases an oscillatory, multidimensional flame located near a side boundary is interacting with one dimensional core flow oscillations. While it is recognized that the characteristics of the investigated diffusion flames are considerably different from those of actual solid propellant flames, it is nevertheless believed that as was the case with the above discussed premixed flames s t ~ d i e s ~ ~ , the findings of this study will improve existing understanding of the mechanisms which control the driving processes in axially unstable solid propellant rocket motors. W

Erosive Augmentation of Solid Propellant Burning Rate: Motor Size Scaling Effect

Abstract: Two different independent variable forms, a difference form and a ratio form, were investigated for correlating the normalized magnitude of the measured erosive burning rate augmentation above the threshold in terms of the amount that the driving parameter (mass flux or Reynolds number) exceeds the threshold value for erosive augmentation at the test condition. The latter was calculated from the previously determined threshold correlation. Either variable form provided a correlation for each of the two motor size data bases individually. However, the data showed a motor size effect, supporting the general observation that the magnitude of erosive burning rate augmentation is reduced for larger rocket motors. For both independent variable forms, the required motor size scaling was attained by including the motor port radius raised to a power in the independent parameter. A boundary layer theory analysis confirmed the experimental finding, but showed that the magnitude of the scale effect is itself dependent upon scale, tending to diminish with increasing motor size.

Fiber optic solid/hybrid rocket motor grain regression rate sensor

Abstract: In one embodiment a bundle of optic fibers of different lengths are inserted through the case wall of a solid fuel or a hybrid rocket motor and extends through the wall of the fuel grain to the edge of the bore in center thereof As the grain burns in a normal manner, from the center toward the outer periphery of the case, the bright flame of ignition is seen at the end of the fiber optic fiber adjacent to the flame The length of each fiber in the bundle is used to determine the burn rate and amount of grain left to ignite In another embodiment different length fibers of a bundle are looped into the grain and back to the exterior of the case and light emitting diodes are positioned at one end of each fiber with a light-detecting means at the other end The light through the diodes is monitored to detect a no-light condition which exists when a fiber is destroyed by the burn In yet another embodiment a single optic fiber is used with a light source and a light reception means at one end As the end of the fiber is destroyed by the ignition of the grain, the time it takes the light from the light source to return to the light sensor is used to establish the burn rate and amount of remaining fuel grain at any given time during the burn

A semi-analytical model for pressure-time history of granular pyrotechnic materials in a closed system

Abstract: Recent experiments with Titanium Subhydride Potassium Perchlorate burning under a zero-volume firing condition have shown that compaction occurs ahead of the burn front. A multiphase flow model can be used to obtain the pressure-time history. However, the number of assumptions on parameters and models must be made to obtain a solution. In this work a semi-analytical model is developed to determine the pressure-time history based on the measured burning front velocity. There are experimental evidence that the burning velocity is a slowly varying function of compaction. It is assumed that the burning particles have a given size distribution and they burn parallel to their outer surface. The burning rate is assumed to be pressure dependent. The penetration of gaseous combustion products into the pores ahead of the burn front and the compaction of the charge can be included by introducing suitable pressure dependent parameters. The formulation of the problem results in an integro-differential equation for the pressure-time history of gaseous combustion products in the close system. This equation is solved numerically to obtain the pressure. A parametric study is performed to find the sensitivity of pressure to different parameters. 15 refs., 5 figs.

Burn Rate Mechanisms. Volume 2. A Model for Ammonium Nitrate Propellant Combustion

Abstract: : When the space shuttle is launched, each Solid Rocket Motor (SRM) booster generates approximately 100 tons of HCI in the exhaust. Thus, with the two boosters, over 200 tons of HCI is generated from each space shuttle launch. The HCI presents several potential problems. The HOI provides nucleation sites for moisture in the air - with the result being the condensed, hydrated form of hydrochloric acid (HCI.H2O). The resultant acid cloud results in a very visible cloud which through further condensation can result in acid rain. The acid rain is obviously detrimental to both plant and animal life dependent on the concentration levels it achieves as it reaches the ground. In addition, at higher elevations, the HOI can react with the protective layer of ozone causing a degradation of some of the ozone. The HO comes from the AR (ammonium perchlorate) which is the oxidizer in the solid propellant for the 5PM boosters. The SRM propellant contains approximately 70% AP which generates slightly more than 21 weight percent HG in the exhaust gases. For these reasons it is desirable to find a replacement for the AP to eliminate the HCI in the rocket exhaust.

Simplified Burn-Rate Model for CMDB Propellants

Abstract: A single model has been proposed to predict the burning rates of bimodal AP,RDX and aluminum containing CMDB propellants. This is done in terms of the respective physical constants on the basis of a recently developed model of combustion of CMDB propellants. The study has been carried out to examine the effects of changes in propellants composition, AP particle size and pressures on burning rate. Computer programs were developed for this purpose and the results obtained for typical sets of input data have been presented and compared with the actual results.