Burning Rate Augmentation Research Articles

Choice of reaction progress variable under preferential diffusion effects in turbulent syngas combustion based on detailed chemistry direct numerical simulations

The combustion of hydrogen and carbon-monoxide mixtures, so-called syngas, plays an increasingly important role in the safety context of non-fossil energy generation, more specifically in the risk management of incidents in process engineering plants for ammonia synthesis and in nuclear power plants. In order to characterize and simulate syngas/air combustion on industrially relevant scales, subgrid modelling is required, which is often based on a reaction progress variable. To understand the influence of different fuel compositions, turbulence intensities and flame topologies on different possible definitions of reaction progress variable, detailed chemistry direct numerical simulations data of premixed, lean hydrogen/air and syngas/air flames has been considered. A reaction progress variable based on normalized molecular oxygen mass fraction has been found not to capture the augmentation of the normalized burning rate per unit flame surface area in comparison to the corresponding 1D unstretched premixed flame due to preferential diffusion effects. By contrast, reaction progress variables based on other individual species, such as hydrogen, can capture the augmentation of the rate of burning well, but exhibit a pronounced sensitivity to preferential diffusion effects, especially in response to flame curvatures. However, a reaction progress variable based on the linear combination of the main products can accurately represent the temperature evolution of the flame for different mixtures, turbulence intensities and varying local flame topology, while effectively capturing the augmentation of burning rate due to preferential diffusion effects. However, its tendency to assume values larger than 1.0 in the regions of super-adiabatic temperatures poses challenges for future modeling approaches, whereas the reaction progress variable based on hydrogen mass fraction remains bound between 0.0 and 1.0 despite showing deviations in comparison to corresponding variations obtained from the unstretched laminar flame depending on flame curvature variations.

Effect of Acoustic Pressure Oscillations on Burning Rate Augmentation of Composite Solid Propellants at Different Initial Grain Temperatures

Effect of Acoustic Pressure Oscillations on Burning Rate Augmentation of Composite Solid Propellants at Different Initial Grain Temperatures

Burning rate augmentation in solid fuel ramjets by swirling inlet air stream

The influence of the inlet swirling flow on the regression rate of the fuel in the combustion chamber of the solid fuel ramjet is investigated in this study using numerical simulations. The finite-volume solver of the compressible turbulent reacting flow is developed to study the flow field where the burning rate is computed using the conjugate heat transfer method for the solid fuel. The correlation is found for the maximum regression rate versus an imposed inlet swirl when the linear distribution of the circumferential velocity is applied at the inlet stream. Although the regression rate augmentation is considerable due to the swirling flow field in the combustor, it is shown that the swirl is effective if is applied near the shear layer of the backstep flow in the combustor. The modified swirler with short blades is suggested to be used in solid fuel ramjets to increase the regression rate of the fuel and improve the performance, but with lower pressure loss.

A New Erosive Burning Model of Solid Propellant Based on Heat Transfer Equilibrium at Propellant Surface

Erosive burning refers to the augmentation of propellant burning rate appears when the velocity of combustion gas flowing parallel to the propellant surface is relatively high. Erosive burning can influence the total burning rate of propellant and performance of solid rocket motors dramatically. There have been many different models to evaluate erosive burning rate for now. Yet, due to the complication processes involving in propellant and solid rocket motor combustion, unknown constants often exist in these models. To use these models, trial-and-error procedure must be implemented to determine the unknown constants firstly. This makes many models difficult to estimate erosive burning before plenty of experiments. In this paper, a new erosive burning rate model is proposed based on the assumption that the erosive burning rate is proportional to the heat flux at the propellant surface. With entrance effect, roughness, and transpiration considered, convective heat transfer coefficient correlation proposed in recent years is used to compute the heat flux. This allows the release of unknown constants, making the model universal and easy to implement. The computational data of the model are compared with different experimental and computational data from different models. Results show that good accuracy (10%) with experiments can be achieved by this model. It is concluded that the present model could be used universally for erosive burning rate evaluation of propellant and performance prediction of solid rocket motor as well.

Acoustic pressure oscillation effects on mean burning rates of plateau propellants

• Influence of acoustic pressure oscillations on mean burning rates of plateau propellants. • Non-aluminized and aluminized propellants exhibit plateau burning trends are examined with corresponding matrixes. • Magnitudes of imposed acoustic amplitudes are varied from 0.1% to 1.4% with respect to mean chamber pressure conditions. • Both non - aluminized propellants and aluminized propellants show significant enhancement in the mean burning rate due to the fluctuations imposed by acoustic pressure amplitudes and frequencies on the propellant combustion. • Maximum burning rate augmentation factors of non-aluminized and aluminized are 1.27 and 1.47 respectively. Combustion instabilities constitute a well pronounced problem in all large rocket motors due to the inherent acoustic pressure oscillations established based on the port geometries. Suppression of the combustion instabilities in the solid rocket motor is required for controlling the mean burning rate variation which arises due to the interaction of acoustic pressure wave with propellant combustion. An experimental study has been carried out to investigate the effects of acoustic pressure oscillations on mean burning rates of non-aluminized and aluminized propellants which exhibit low pressure exponent index (n) in the burning rate trends. Steady and unsteady mean burning rates are determined from combustion photography method using a window bomb test facility over the pressure range of 17 MPa. A rotary valve is coupled with the window bomb setup to generate acoustic pressure oscillations inside the test chamber (cylindrical pressure vessel), which imposes the required frequencies of 140, 180 and 240 Hz respectively. The acoustic pressure amplitudes are varied from 0.04% to 1.4% of the mean chamber pressure. Both non-aluminized propellants and aluminized propellants have shown significant enhancement in the mean burning rate due to the fluctuations imposed by acoustic pressure amplitudes and frequencies on the propellant combustion. The enhancement in the mean burning rate also depends upon dynamic response of the flame to the excited frequencies and acoustic pressure amplitudes. The plateau burning behaviour of the non-aluminized propellant is completely distorted whereas it is retained in aluminized compositions. Conversely, it also shifts the mean pressure range of plateau burning rate trend. The maximum burning rate augmentation factors resulted from imposed acoustic pressure wave on non-aluminized and aluminized propellants are observed as 1.27 and 1.47 respectively.

Measurement of admittance and acoustic augmentation of burning rate of composite solid propellants using Laser Doppler Velocimetry

The present work involves the use of Laser Doppler Velocimetry (LDV) technique to measure the acoustic admittance of aluminized composite solid propellant (AP/HTPB/Al) by experimentally measuring the instantaneous velocities of the particles of solid propellant during burning while the acoustic pressure oscillations are being imposed. Both velocity fluctuations and pressure fluctuations are measured synchronously. The probe volume of the laser is maintained at a constant distance from the burning surface by using a linear actuator. Tests were conducted for two aluminized composite solid propellants at low frequency ranges (130-312 Hz) of the longitudinal acoustic modes and mean chamber pressures ranging from 1 to 4 MPa. A rotary valve is designed to acoustically excite the chamber to the required resonant frequency. It was found that admittance rises with frequency and peaks around 260 Hz and then falls. The use of linear actuator mechanism significantly improves the velocity data rate (number of data points). Maintaining the probe volume at 800 to 900 µm above the propellant surface is very crucial to determine the admittance accurately. Test results revealed that a good agreement between the steady state mean burning rates with the window bomb experiments is achieved. The measured mean burn rates under imposed acoustic pressure oscillations shows the augmentation up to 55%. The admittance measured for the metallized composite propellant using this technique and the comparison with T-burner shows good agreement. The comparison of responses with LDV and T-burner and the augmentation factor due to acoustic oscillations using LDV are reported for the first time.

Mean Burning Rate Variation in Composite Propellant Combustion Due to Longitudinal Acoustic Oscillations

This paper focuses on investigating “combustion instability,” a phenomenon that mainly involves the interaction of the propellant flames with the acoustic oscillations prevalent in full-scale rocket motors. The effect of excited acoustic pressure oscillations on the mean burning rate of solid propellants was estimated over the pressure ranges from 1 to 7 MPa using a window bomb facility coupled with a rotary valve. Both non-aluminized and aluminized composite propellants were used. A rotary valve was used to drive the acoustic oscillations at frequencies of 140, 180 and 240 Hz with the pressure amplitude perturbations ranging from 0.1 to 1.4% of mean chamber pressure. Frequency shift due to propellant combustion was also investigated for both the types of propellants. The acoustic oscillations enhance the heat transfer between the flame and propellant burning surface, altering the mean burning rate. The acoustic pressure oscillations influence the dynamics of aluminum particles and its agglomeration processes, which modifies the mean burning rate. The variations in excited frequencies show a significant impact on the mean burning rate. The burning rate augmentation factor shows that acoustic excitation is more predominant at low pressures and high frequencies, whereas it is relatively marginal at high pressures. The evaluated maximum augmentation factors are 1.45 and 1.51 for nonaluminized and aluminized propellants when compared with those without oscillations.

Parameters Affecting the Erosive Burning of Solid Rocket Motor

Increasing the velocity of gases inside solid rocket motors with low port-to-throat area ratios, leading to increased occurrence and severity of burning rate augmentation due to flow of propellant products across burning propellant surfaces (erosive burning), erosive burning of high energy composite propellant was investigated to supply rocket motor design criteria and to supplement knowledge of combustion phenomena, pressure, burning rate and high velocity of gases all of these are parameters affect on erosive burning. Investigate the phenomena of the erosive burning by using the 2’inch rocket motor and modified one. Different tests applied to fulfil all the parameters that calculated out from the experiments and by studying the pressure time curve and erosive burning phenomena.

Regression Rates Study of Mixed Hybrid Propellants

The low regression rates of classic rocket fuels lead to large internal ports that limit potential applications. This experimental study investigated the increase in regression rate that results from adding a solid oxidizer and a catalyst to a fuel grain. The configuration is named a hybrid to signify solid oxidizer and catalyst in the grain. A design of experiments approach guided fuel formulation to systematically control levels of ammonium perehlorate from 25% to 30%, ferric oxide from 0 to 5%, and hydroxyl-terminated polybutadiene from 70% to 75%. The 1.5-in. diam. port, 12-in. long center perforated grains were burned with gaseous oxygen at pressure levels from 150 to 550 psig and port flux levels from 0.1 to 0.4 lbm/s-in. 2 . The results show that the mixed propellants burn as a function of both pressure and mass flux. A grain formulation having 27.5 % ammonium perchlorate and 2.5% ferric oxide provided the maximum burning rate augmentation (447%) among the formulations tested.

Star-grain rocket motor – nonsteady internal ballistics

The nonsteady internal ballistics of a star-grain solid-propellant rocket motor are investigated through a numerical simulation model that incorporates both the internal flow and surrounding structure. The effects of structural vibration on burning rate augmentation and wave development in nonsteady operation are demonstrated. The amount of damping plays a role in influencing the predicted axial combustion instability symptoms of the motor. The variation in oscillation frequencies about a given star grain section periphery, and along the grain with different levels of burnback, also influences the means by which the local acceleration drives the combustion and flow behaviour.

Statistical modeling of thermal radiation transfer in buoyant turbulent diffusion flames

A statistical (Monte Carlo) method for radiative heat transfer has been incorporated in CFD modeling of buoyant turbulent diffusion flames in stagnant air and in a cross-wind. The model and the computational tool have been developed and applied to simulate both burner flames with controlled fuel supply rate and in self-sustained pool fires with burning rates coupled with flame radiation. The gas–soot mixture was treated either as gray (using the effective absorption coefficient derived from total emissivity data or the Planck mean absorption coefficient) or as non-gray (using the weighed sum of gray gases model). The comparison of predicted radiative heat fluxes indicates applicability of the gray media assumption in modeling of thermal radiation in case of high soot content. The effect of turbulence-radiation interaction is approximately taken into account in calculation of radiation emission, which is corrected to allow for temperature self-correlation and absorption-temperature correlation. In modeling buoyant propane flames in still air above 0.3 m diameter burner, extensive comparison is presented of the predictions with the measurements of gas species concentrations, temperature, velocity and their turbulent fluctuations, and radiative heat fluxes obtained in flames with different heat release rates. Similar to previously published experimental data, the predicted burning rate of flames above the acetone pools exposed to flame radiation increases with the pool diameter and approaches a constant level for large pool sizes. The magnitude of predicted burning rates is shown to be in agreement with the reported measurements. Augmentation of burning rate of the pool fire in a cross-wind because of increased net radiative heat flux received by the fuel surface and non-monotonic dependence of burning rate on cross-wind velocity, subject to the pool diameter, is predicted. The statistical treatment of thermal radiation transfer has been found to be robust and computationally efficient.

Chemical characterization of plasma-treated propellant samples

The morphological and chemical characterization of M30 propellant grains recovered after conventional or plasma ignition in interrupted closed bomb experiments has been performed. The capillary material used during plasma ignition was either polyethylene (PE) or polyethylene terephthalate (Mylar). Propellants were extinguished at pressures between 35 and 100 MPa. The samples are of interest due to previous reports of apparent burning rate augmentation with plasma ignition of M30, compared to conventional ignition, in non-interrupted closed bomb experiments [1]. Birk [2] has reported that burning rate augmentation appears to occur during the plasma event, but there was no evidence for post-plasma augmentation. Birk has also analyzed vivacity curves from plasmaignited M30 and the data indicated that the grains appear to burn regressively, suggesting an increase in surface area was generated beyond that expected from normal grain burning [2]. In the current work, differences between extinguished grains from conventional and plasma ignition were primarily physical or morphological in nature. For the extinguished grains from PE plasma ignition there appeared to be a stripping of NQ crystals in the perforations. Along the outer surfaces of the grain ignited with plasma, hot embedded particles and increased surface area were observed, and the melt layer was immeasurably thin. Although extensive chemical characterization has been performed, there appears to be very little chemical difference between the burned surfaces of the plasma and conventionally ignited samples. The SEM analysis of the perforations showed that NQ depletion in the perforations was most evident at the lowest blowout pressure used (35 MPa) and seemed to diminish at higher pressure, apparently becoming obscured by “normal” burning processes. Thus, it appears that plasma-propellant interactions and associated effects occur only very early during ignition. This is consistent with the fact that few differences in the chemical composition are observed for plasma and conventionally ignited samples. A depletion of NQ in the perforations may contribute to a true burning rate augmentation through chemical reactions. However, a true chemical burning rate augmentation as the only mechanism for the M30 non-interrupted behavior would not be consistent with regressive burning.

Analysis of burning rate phenomena and extinguished solid propellants from an interrupted closed bomb with plasma igniter [ETC launchers

The morphological and chemical characterization of M30 propellants recovered after conventional and plasma ignition with polyethylene (PE) and polyethylene terephthalate (Mylar) in an interrupted closed bomb (extinguished at pressures between 35 and 100 MPa) has been performed. This work was undertaken to investigate mechanisms associated with the apparent burning rate augmentation of M30 under conditions of plasma ignition compared to conventional ignition. Burning rate augmentation of solid propellants with plasma ignition was found to occur during, but not after, the plasma event. The following factors are evidence of a more forceful process for plasma ignition compared to conventional ignition: a) the initial pressure rise was greater with plasma ignition; b) metals from the closed bomb hardware, and stripping of NQ from M30 propellants was observed in the perforations; c) irregular surface burning (possibly due to ablation). And hot embedded particles have been observed in grains after plasma, but not conventional, ignition. Extensive chemical analyses of extinguished grains have been performed, but no significant differences were revealed between plasma and conventional ignition, quite possibly because any early chemical interactions are obscured by normal burning processes occurring later.

Plateau Burning Characteristics of AP Based Azide Composite Propellants

AbstractThe site and mechanism by which iron oxide catalyst acted to enhance burning rate and produced plateau burning behavior at high pressure was studied. The condensed phase chemistry study was conducted by isothermal thermogravimetric analysis (TGA) and differential scanning calorimetry (DSC), and rapid‐scan FTIR spectroscopic technique. Uncatalyzed ammonium perchlorate (AP) based azide composite propellant showed unstable combustion at relatively lower pressure region. The heat balance at the buring surface would be unstable at these pressures. However, iron oxide altered the burning property of the propellant and enhanced the burning rate with the plateau‐mesa burning characteristics. Such pressure insensitiveness of the burning rate indicated that the condensed phase chemistry played important role in the catalytic mechanism of action. According to the microrocket motor tests, physical effect, melted fuel binder covered the AP particles and prevented the further decomposition of AP, had not affected the plateau burning. Fe2O3 was more effective on the burning rate augmentation than Fe3O4. However, the pressure exponent of the burning rate point of view Fe3O4 was favored catalyst to the propellant used here.

Burning Rate Augmentation of BAMO Based Propellants

AbstractThermal decomposition and the burning properties of BAMO based propellants with HMX or AN/HMX have been investigated. The heat generated by the azide binder decomposition initiated and accelerated the thermal decomposition of HMX and AN. Ammonium perchlorate (AP) and lead stearate with carbon black significantly altered the mechanisms of the thermal decomposition and the burning properties of the HMX based propellants. AP showed an increase in burning rate with a slight decrease in burning rate pressure exponent. The lead catalyst yielded high value of the burning rate with the lowest pressure exponent. The ammonium dichromate also influenced the mechanisms of the thermal decomposition and the burning properties of the AN/HMX samples. The combination of ammonium dichromate and copper chromite was the most effective on the burning rate augmentation of AN/HMX based propellants. AN sublimed and evaporated from the condensed phase and mainly reacted exothermically in the gas phase HMX and AN/HMX based propellants showed smokeless burning characteristics in the small rocket motor combustion tests.

Erosive burning of solid propellants

Presented here is a review of the experimental and modeling work concerning erosive burning of solid propellants (augmentation of burning rate by flow of product gases across a burning surface). A brief introduction describes the motor design problems caused by this phenomenon, particularly for low port/throat area ratio motors and nozzleless motors. Various experimental techniques for measuring crossflow sensitivity of solid propellant burning rates are described, with the conclusion that accurate simulation of the flow, including upstream flow development, in actual motors is important since the degree of erosive burning depends not only on local mean crossflow velocity and propellant nature, but also upon this upstream development. In the modeling area, a brief review of simplified models and correlating equations is presented, followed by a description of more complex numerical analysis models. Both composite and double-base propellant models are reviewed. A second generation composite model is shown to give good agreement with data obtained in a series of tests in which composite propellant composition and heterogeneity (particle size distribution) were systematically varied. Finally, the use of numerical models for the development of erosive burning correlations is described, and a brief discussion of scaling is presented.

The effect of cross-flow temperature on the burning rate in a solid propellant

It is well-known that the burning rate of the propellant for a solid rocket motor or a solid fuel ramjet increases as the cross-flow velocity increases. It has been shown, however, in many previous studies that the cross-flow temperature does not have a strong effect on the burning rate. In this study, the effect of the cross-flow temperature, on the burning rate was examined by using a double-base propellant and a small rocket motor. The augmentation of burning rate with cross-flow was found to have a weak dependence on the cross-flow temperature for temperatures above 1500 K. However, for cross-flow temperatures under 1500 K, the burning rate decreased significantly as the temperature decreased. Namely, a threshold temperature was found to exist in the effect of the cross-flow temperature. A simple model based on the energy and mass transfer between the cross-flow and the fizz zone was used to examine the effect of the cross-flow on the heat feedback from the gas phase. In this model, two parameters Φ c and Φ k were used. The parameter Φ c is related to the change of the fizz zone temperature due to turbulence in the fizz zone. The parameter Φ k is related to fizz zone physical properties such as thermal conductivity or flame geometry. The calculated results using these parameters had good agreement with the experimental data. The model showed that the parameter Φ c was important in determining the burning rate. Namely, the energy transfer between the cross-flow and the fizz zone due to turbulence increased the fizz zone temperature and the reaction rate in the fizz zone. As a result, the heat feedback from the gas phase increased and the burning rate also increased significantly.

Role of Boron in Burning Rate Augmentation of AP composite propellants

AbstractEffect of the addition of boron particles on the burning rate of solid propellants was examined. The propellants tested in this study consisted of ammonium perchlorate (AP) as an oxidizer and carboxyl terminated polybutadiene as a fuel binder. The propellant burning rate is increased significantly by the addition of a small amount of boron particles. The burning rate augmentation is dependent largely on the size and concentration of the boron particles mixed. Thermochemical experiments revealed that the boron particles react with the decomposed gases of AP on and just above the propellant burning surface. The heat flux transferred back from the gas phase to the burning surface of the propellant increases with increasing the total surface of the boron particles mixed within the unit mass of propellant. The burning rate augmentation is correlated to the heat of reaction generated by the oxidation reaction of boron particles.

NORMAL ACCELERATION MODEL FOR COMPOSITE-PROPELLANT COMBUSTION

A new engineering model is presented for the description of normal or radial acceleration effects on the combustion of composite solid propellants having a low metal content. This phenomenological, one-dimensional, flame-sheet model is based on the concept of a compressed combustion zone due to the retarding effects of a normal acceleration field. The resulting equations in their general form are free of experimentally derived coefficients. A comparison of predictions and experimental data from several sources indicates fair agreement with regard to the burning rate augmentation trends at lower acceleration field strengths.

Low Burning Rate Aluminized Propellants in Acceleration Fields

To reduce slag formation and burning rate increases, the effects of radial acceleration (up to 20 g) and low pressure (2-7 MPa) on the combustion of low-burning rate aluminized propellants (87 formulations) were investigated. Experiments were performed to obtain agglomerate size distribution, slag formation, and burning rate data. Data sources included photographs under cross flow and acceleration conditions, scanning electron probe (Cl and Al) images, particle collection combustors, and rocket motors. The tendency to form slag increases as the agglomeration size at the burning surface increases. A lower burning rate increases acceleration effects. However, residue formation is more sensitive to formulation than is burning rate; for example, bimodal vs trimodal AP can cause significant changes in residue in the absence of burn rate changes; RDX produces larger agglomerates than HMX. Through understanding obtained from this research, reduction in agglomeration, slag formation, and burning rate augmentation are directly attributable to systematic changes in formulation ingredient levels and sizes.