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Burn rate (chemistry)

About: Burn rate (chemistry) is a research topic. Over the lifetime, 847 publications have been published within this topic receiving 8908 citations. The topic is also known as: Burning rate.


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Journal ArticleDOI
TL;DR: In this paper, the authors compared the error estimates generated by NASTRAN and those computed with Eq. (8), especially for the higher modes, are due to terms O(kL) neglected in the kinetic energy error estimation.
Abstract: estimates, Eqs. (8) and (9) and those obtained from NASTRAN calculations, are compared in Table 2. The large differences between the error estimates generated by NASTRAN and those computed with Eq. (8), especially for the higher modes, are due to terms O(kL) neglected in the kinetic energy error estimation. These terms become significant as (ir/N) 4 approaches unity, that is, as the number of elements per half wavelength decreases.

5 citations

Patent
26 Oct 1998
TL;DR: In this article, an energetic solid rocket motor propellant having one or more plateau regions of low operating pressure exponent is disclosed. The propellant is formulated from ingredients including an energetic polyoxetane, an effective amount of a plasticizer, an inorganic oxidizer in at least two discrete particle size ranges, and a refractory oxide burn rate modifier.
Abstract: An energetic solid rocket motor propellant having one or more plateau regions of low operating pressure exponent is disclosed. The propellant is formulated from ingredients including an energetic polyoxetane, an effective amount of a plasticizer, an inorganic oxidizer in at least two discrete particle size ranges, and a refractory oxide burn rate modifier.

5 citations

01 Jan 2006
TL;DR: In this paper, a comparison of small scale hybrid data to that of larger scale data indicates that the fuel burn rate goes down with increasing port size, even with the same oxidizer flux.
Abstract: Hybrid rocket motors can be successfully demonstrated at a small scale virtually anywhere. There have been many suitcase sized portable test stands assembled for demonstration of hybrids. They show the safety of hybrid rockets to the audiences. These small show motors and small laboratory scale motors can give comparative burn rate data for development of different fuel/oxidizer combinations, however questions that are always asked when hybrids are mentioned for large scale applications are - how do they scale and has it been shown in a large motor? To answer those questions, large scale motor testing is required to verify the hybrid motor at its true size. The necessity to conduct large-scale hybrid rocket motor tests to validate the burn rate from the small motors to application size has been documented in several place^'^^.^. Comparison of small scale hybrid data to that of larger scale data indicates that the fuel burn rate goes down with increasing port size, even with the same oxidizer flux. This trend holds for conventional hybrid motors with forward oxidizer injection and HTPB based fuels. While the reason this is occurring would make a great paper or study or thesis, it is not thoroughly understood at this time. Potential causes include the fact that since hybrid combustion is boundary layer driven, the larger port sizes reduce the interaction (radiation, mixing and heat transfer) from the core region of the port. This chapter focuses on some of the large, prototype sized testing of hybrid motors. The largest motors tested have been AMROC s 250K-lbf thrust motor at Edwards Air Force Base and the Hybrid Propulsion Demonstration Program s 250K-lbf thrust motor at Stennis Space Center. Numerous smaller tests were performed to support the burn rate, stability and scaling concepts that went into the development of those large motors.

5 citations

Proceedings ArticleDOI
08 Jan 1990
TL;DR: In this paper, the interaction between diffusion flames, stabilized on the side wall of a long rectangular duct, and an axial acoustic field was investigated experimentally, and 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.
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

5 citations

Proceedings ArticleDOI
06 Jan 1997
TL;DR: In this paper, a pressure insensitive force transducer obtained from Zarko was used in the laser recoil device at pressures from 1 to 6 atm, and the laser flux was measured for HMX, RDX and N5 up to 60 W/cm2 at 13.7 and 60 psia.
Abstract: HMX and RDX pellets were pressed and tested in the laser recoil device at pressures from 1 to 6 atm. The laser flux had an average power of 31 W/cm2, with sinusoidal oscillations from 15 to 43 W/cm2. A pressure insensitive force transducer obtained from Zarko was used in these tests. The force transducer has a resonant frequency at 160 Hz, but performs well at frequencies below 100 Hz. Samples inhibited with halocarbon grease, and with glass tubing were compared. HMX showed twice the response amplitude of RDX. Both HMX and RDX have a broad flat response. In contrast with the previously measured N5 double base propellant, the recoil response of neither HMX nor RDX changed significantly with pressure up to 60 psia. Burn rate vs. laser flux was measured for HMX, RDX and N5 up to 60 W/cm2 at 13.7 and 60 psia. At one atmosphere the rate vs. flux slope of HMX and RDX is about the same, but N5 is 3 times. Other ingredients tested include GAP, GAP/BTTN, NMMO gumstocks, BAMO pellets, and IRQ minimum smoke propellant. GAP and BAMO would not burn with a stable flame, but chuffed large puffs of nitrogen. NMMO ignited readily, but did not produce force oscillations. GAP/BTTN had a very high unstable laser recoil response, even bigger than N5 double base propellant, which did not decrease significantly with pressure.

5 citations


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Performance
Metrics
No. of papers in the topic in previous years
YearPapers
202310
202220
202116
202015
201918
201811