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.

Strand Burner Results of AFP-001 Propellant with Inert Coating for Temperature Compensation

Abstract: : The US Army Research Laboratory is currently exploring methods of manufacturing propellants whose performance is negligibly impacted by the wide range of operating temperatures demanded of a fielded propellant. The current approach is to coat the propellant with an inert material such that the amount of exposed base grain available to initially burn is controlled through the temperature-dependent mechanical properties of the coating. In effect the coating inhibits surface area initially available to burn at hot temperatures and exposes the greatest amount of base grain at cold temperatures. The end result is a propellant performance (i.e., burning rate) profile that is independent of initial temperature. This report discusses the strand burner results of 3 chosen candidate coatings that were applied to an AFP-001 base grain propellant. Included will be the base grain's burning rate at a pressure of up to 10 MPa and the burning rates of the coated propellants, also up to 10 MPa.

A computer-based tool for preliminary design and performance assessment of Continuous Detonation Wave Engines

Abstract: As classical liquid rocket propulsion systems seem to have reached their limit in terms of performance, the promising potential of Continuous Detonation Wave Engines (CDWE) has led to a growing interest in its use for space applications. In theory, the detonation regime of combustion offers a promising alternative for traditional fuel burning methods based on deflagrations. The detonation process is characterised by a higher burn rate and energy release rate, which produces an extreme pressure and temperature rise. It is therefore more similar to a constant-volume process than a constant-pressure process typical of conventional combustion. Therefore, detonation engines are expected to have a higher thermodynamic efficiency and specific impulse than conventional Liquid Rocket Engines (LRE). Moreover, due to the short combustion zone, the length of the combustion chamber may be shortened and the engine mass lowered. The principle of CDWE is based on the formation of rotating detonation waves in an annular cylindrical combustion chamber. The propellants are supplied from one side and combustion products expand towards the other side of the chamber, and reach supersonic velocity without the need for a geometrical throat. Compared to the Pulsed Detonation Engine (PDE), CDWE can provide a nearly steady thrust level (higher thrust-to-weight ratio) and requires only one detonation initiation. It also generates a reduced vibrational environment and is more suitable for operation in a low pressure environment. Nevertheless, the CDWE includes many theoretical and practical challenges of its own that need to be understood and overcome. To study the theoretical performance of CDWE and its impact on launch vehicles, a simple computer-based tool has been developed which allows fast preliminary design and performance assessment of the engine. The tool comprises an engineering model based on theoretical and empirical relations from literature, and provides an initial estimation of the engine performance, dimensions and mass for LOX/LH2 combination only. Results have been verified with published data from previous numerical and experimental studies. The tool has then been used for design case studies to investigate the hypothetical integration of CDWE in the Ariane 5 ME (A5ME) launch vehicle and to assess the corresponding achievable performance gain. This is done considering the direct exchange of the current cryogenic upper and main stage engines (Vinci and Vulcain 2) with an equivalent CDWE designed to operate at similar injection pressure, mixture ratio and mass flow rate, allowing most of the configuration of tanks and turbopumps to remain the same (in case of the core stage). Taking into account the preliminary CDWE designs, trajectory optimizations have been performed to determine the resulting payload performance gain. The results showed that an increase in payload performance can be achieved due to the higher specific impulse and the reduced mass of CDWE compared to LRE. With CDWE in the core or upper stage only, the payload performance can be increased by roughly 6 and 7% respectively, and with CDWE in both main and upper stages this is about 14%. These results clearly demonstrate the large theoretical potential of CDWE for rocket propulsion applications. Nevertheless, it is important to consider the impact of the idealized and simplified assumptions made in the model on the results. Of large influence are the assumed initial conditions, in particular the injector pressure loss and the fresh mixture Mach number, as they directly influence the calculated detonation parameters. Also the injection temperature strongly impacts the detonation pressure ratio, especially at low (cryogenic) temperatures. Furthermore, the assumption of ideal Chapman-Jouguet (CJ) detonation leads to higher detonation pressures and velocities than in reality. Overall, the model considers an ideal engine which implies steady-state engine operation, and neglects heat losses and friction at the walls. Additional phenomena causing pressure losses such as the expansion of propellant jets, non-uniform propellant mixing, contact surface burning, shocks and shear layers in the flow expansion zones, etc. are not considered in the model, although they may have a significant impact on the CDWE performance in reality. For a more reliable analysis, the tool could be improved by replacing the current CJ detonation calculation with a more accurate detonation model, such as the Zel’dovich, Von Neumann, and Doring (ZND) model. It would require the implementation of a detailed chemistry model, which allows the calculation of dynamic detonation parameters as well as an equilibrium approach for the flow expansion through the chamber and the nozzle. This would fully eliminate the need for the CEA program to determine the detonation properties. Furthermore, a more accurate mass estimation could be obtained with the availability of additional data on CDWE materials, wall thicknesses, component masses, heat fluxes, etc. which could not be found so far. For future studies with the design tool it would be interesting to optimize CDWE for different injection pressures and mixture ratios. Since similar chamber pressures can be obtained with CDWE at lower injection pressures than for classic LRE, this could lead to a significant relief of the demand on the injection system. CDWE also has the potential to obtain a similar performance as LRE at higher mixture ratios, which would allow the use of smaller tanks for the same propellant mass in the case of LOX/LH2.

The Research on Transient Burning Rate of Solid Propellant by Digital Image Processing

Abstract: In order to obtain the burn rate of the solid propellant that is the important parameter of transient burning, the new method named digital image processing is presented. In the article , the principle of digital image processing is analysed; The burning face of the sample in the each time is located according the image and the coordinates of the burning face is obtained. In experiment the transient burn rate is measured by digital image processing and the accuracy is acceptable.

The Influence of Propellant Loading Density on Computed Burn Rate in a Mini-Closed Bomb

Abstract: : An investigation was conducted to determine what loading density should be used to calculate propellant thermochemical properties used in closed-chamber data analysis to minimize the differences in computed burn rates observed as the propellant loading density in the closed chamber varies. A comparison between the traditional loading density of 0.2 g/sq cm and the actual propellant loading density was made. The traditional r = bP(exp n) burn rate law was used as the basis for the comparison.

Numerical Simulation for Combustion Characteristic of Polyethylene in Solid Fuel Ramjet

Abstract: To accurately predict the combustion characteristic of polyethylene in solid fuel ramjet,the viscosity and thermal conductivity coefficients of C2H4,CO,CO2,H2O,O2 and N2 were estimated at different temperature.And mass,momentum and energy sources were added using UDF.The results show that the trend of regression rate against the change of the total temperature or mass flow rate of air is agreed well with existing experiments.The maximum errors of mean regression rates and aft-mixing combustion chamber temperatures are 11.3% and 5.2%,respectively.And the trend of aft-mixing combustion chamber temperature is opposite with the mass flow rate of air.When the total temperature of air increases or the mass flux decreases,the length of aft-mixing combustion chamber needs to be longer to fully burn the fuel.With geometric similarity,when other operation conditions are the same,the heat flux from flow field to surface of propellant and regression rate are greater in smaller size SFRJ.