The measurement of the regression rate of a hybrid rocket fuel is the subject matter of this paper. This is an involved problem because both the regression rate and the oxidizer mass flux Gox vary with the port diameter, which itself varies during combustion. This paper addresses the issue of how Gox should be varied to obtain the regression rate correctly. The most commonly used weight loss method has been employed here. Three methods used here to vary the Gox are artificially changing the port diameter, changing the mass flow rate of the oxidizer, and interrupted burn test. From the experiments conducted, it was evident that the interrupted burn test was the best method. These experiments were also carried out for a larger motor and the results compared quite well with established literature. The role of the recirculation zone size near the head end and the overall heat transferred back to the fuel surface in influencing the regression rate have been explained. The role of the recirculation zone diminishes as the length of the rocket motor is increased.

Periyapatna A. Ramakrishna

Department of Aerospace Engineering

FUEL SURFACES

HYBRID FUELS

HYBRID ROCKETS

Mass flow rate

Recirculation zones

REGRESSION RATE

SUBJECT MATTERS

WEIGHT LOSS METHOD

Experiments

Fuels

Rocket engines

Rockets

Regression analysis

IIT Madras is a public technical and research university located in Chennai, Tamil Nadu. Founded in 1959, it is recognised as an Institute of National Importance.

IIT Madras has been ranked as the top engineering institute in India for four years in a row by the National Institutional Ranking Framework of the MHRD

It currently offers undergraduate, postgraduate and research degrees across 16 disciplines in Engineering, Sciences, Humanities and Management. About 596 faculty belonging to science and engineering departments and centres of the Institute are engaged in teaching, research and industrial consultancy.

IIT Madras

Issues Related to the Measurement of Regression Rate of Fast-Burning Hybrid Fuels

Journal of Propulsion and Power

We characterize the effect of punctured discs, that acts as circular protrusions at predefined axial locations of a lab-scale hybrid rocket motor. The fuel employed is paraffin, while the oxidizer is oxygen. The quantities measured are the burn duration and the exhaust plume length. It is seen that the burn duration is reduced for the same quantity of fuel and oxidizer as compared to the case with no protrusion. On the other hand, the exhaust plume length as measured from moderately time-resolved image processing reveals longer plume length. Also observed is that the exhaust plume length for the non-protrusion case reveals sudden dip and flare, which denotes more erratic operation as compared to the case with protrusion. The higher regression rate (low burn time) measured for the case with protrusion is analyzed computationally for isothermal flow. It is seen that higher levels of turbulence as compared to the case of no protrusion aids in pronounced heat and mass transfer resulting in higher regression rate. © 2019 by American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.

AIAA Propulsion and Energy Forum and Exposition, 2019

Study of regression rate characteristics in a hybrid rocket motor

Polytetrafluoroethylene is known to enhance the reactivity of aluminum. The enhancement in the reactivity is produced by mechanically activating the aluminum using polytetrafluoroethylene. In this study, mechanically activated aluminum powder is used along with paraffin wax as the fuel grain for a hybrid propellant. Here, gaseous oxygen is used as oxidizer. To achieve an optimum composition for the best specific impulse, the aluminum fraction in the composition is restricted to 50% of the base fuel composition along with 50% wax/ethylene vinyl acetate (combination of wax and ethylene vinyl acetate, in a ratio of 4∶1). The polytetrafluoroethylene in mechanically activated aluminum powder is added at the expense of wax/ethylene vinyl acetate in the fuel composition. When compared with the composition having pyral aluminum powder (which has flake like particles of very high specific surface area), mechanically activated powder increases the fuel regression rate by ∼40% at an oxidizer mass flux of 35 g∕cm2-s. It is also observed that the combustion efficiency of the hybrid motor is 59% with the use of mechanically activated aluminum powder in comparison to 39% obtained with micron-sized aluminum powder. Copyright © 2018 by Indian Institute of Technology, Madras, India.

Utilization of mechanically activated aluminum in hybrid rockets

It has been reported in literature that the use of protrusion in the combustion chamber of a hybrid rocket motor enhances the regression rate. A detailed study was carried out by using two types of protrusion (a) rubber (which burns along with fuel but at different rate, hence its shape changes with burn time), and (b) graphite (does not burns along with fuel and hence its shape does not changes with burn time). The results showed that the improvement in regression rate with protrusions was only up to a certain fraction of the overall burn time. From the results obtained, it was observed that an X/L of 0.5 was the best location for both types of protrusion. The maximum improvement in regression rate was around 45%. The combustion efficiency was observed to improve from 58% without protrusion to 69% with the use of protrusion, when placed at an X/L of 0.8. It was observed that the use of bluff body and protrusion improves combustion efficiency, when used individually. Hence, combination of the two was used and the combustion efficiency was determined. The combustion efficiency was around 90%, when bluff body was used at the head end and the protrusion at an X/L of 0.8 from head end. © 2014 Elsevier Masson SAS.

Aerospace Science and Technology

Effect of protrusion on the enhancement of regression rate

Iron oxide and copper chromite are the known burn rate enhancers used in a composite solid propellant. Lot of research has been carried out to understand the mechanism or location of action of the burn rate modifiers so as to better tailor the burning rate of a composite propellant. The literature is still very confusing in affirming the mechanism. Here, a systematic study has been carried out, by undertaking experiments at varying levels of combinations of the individual components (ammonium perchlorate, which is oxidizer and hydroxyl terminated poly butadiene, which is both fuel and binder) of composite solid propellant. Firstly, thermal gravimetric analysis, differential scanning calorimetry and burning rate measurements on the individual components are carried out to study the effect of iron oxide and copper chromite on the components themselves. It has been noticed that though both iron oxide and copper chromite are effective on ammonium perchlorate, iron oxide is slightly more effective than copper chromite. Also, copper chromite enhanced the binder melt flow, while iron oxide reduced it. These are followed-up by experiments on sandwich propellants, which give greater insight and enables better understanding of the behavior of iron oxide and copper chromite in composite propellants, as these are simple two-dimensional analogue of the composite solid propellants. Finally, experiments are carried out on the composite solid propellants to obtain a holistic understanding of the behavior/location of action of iron oxide and copper chromite in them. These studies are used to explain certain unexplained but observed phenomena, at the same time elucidating the location of action of these burn rate modifiers in composite solid propellant combustion. Based on these observations, it has been proposed that both iron oxide and copper chromite are primarily acting on the condensed phase. These studies are further complimented with experiments to analyze the thermal conductivity measurements of various propellant samples. This is pursued to understand the reason for the differences in burn rate pressure index for the composite propellants with iron oxide and with copper chromite. It has been understood from these studies that the thermal conductivity of a composite propellant is a key parameter, which affects the burn rate pressure index. Literature has never addressed it from this perspective. © 2014 The Combustion Institute.

Combustion and Flame

Studies on the role of iron oxide and copper chromite in solid propellant combustion

An attempt was made in this paper to determine the regression rate of a hybrid fuel by using combustion chamber pressure. In this method, the choked flow condition at the nozzle throat of the hybrid rocket was used to obtain the mass of fuel burnt and in turn the regression rate. The algorithm used here is better than those reported in the literature as the results obtained were compared with the results obtained using the weight loss method and was demonstrated to be in good agreement with the results obtained using the weight loss method using the same motor and the same fuel and oxidizer combination. In addition, the O/F ratio obtained was in good agreement with those obtained using the weight loss method. The combustion efficiencies obtained were in good agreement with the average values. © 2014 IAA. Published by Elsevier Ltd. All rights reserved.

Acta Astronautica

Measurement of regression rate in hybrid rocket using combustion chamber pressure

In this work, the effect of solid oxidizer (AP) mixed fuel grain (HTPB) on regression rate and performance in GOX hybrid rocket motors is numerically studied. Here the pyrolysis products of AP and the fuel forms a premixed flame very close to the grain surface in addition to the fueloxygen diffusion flame located further away from the grain surface. The presence of premixed flame increases heat feedback to the grain and consequently increases the local regression rate throughout the grain length. This is contrary to the locally limited influence observed while using other regression rate enhancement techniques like swirling oxidizer, or transverse diaphragms/protrusions. Scaling of motors with AP mixed grain of a given composition follows that of a pure fuel grain where the grain geometry and the inlet oxidizer mass flux need to be maintained. © 2015 Elsevier Masson SAS. All rights reserved.

A numerical investigation of oxidizer mixed hybrid rocket motors

The results of a systematic experimental investigation on the methods of enhancing the regression rate in hydroxyl-terminated polybutadiene (HTPB) fuel used in an HTPB/gaseous oxygen hybrid motor are presented. The effects of the addition of ammonium perchlorate (AP) or aluminum in the fuel, the variation of oxidizer-fuel ratio, and the variation of characteristic dimensions of fuel grain are presented. For the extents of the parametric variations considered, while the addition of AP and/or Al and the reduction of grain port diameter enhance the regression rate, the effect due to the latter is the most significant one. Furthermore, the regression rate increases along the axis, and it becomes essentially constant in the port region corresponding to a fuel-rich composition. The possible physical processes for all of these behaviors are discussed. The experimentally obtained exponents of the variables for regression rate (oxidizer mass flux and port diameter) are found to be significantly different from those of the conventional theory. The similarity between the fuel regression rate equation used in solid-fuel ramjet and that obtained in hybrid motor is discussed.

Fuel regression rate in hydroxyl-terminated-polybutadiene/gaseous-oxygen hybrid rocket motors

Combustion aspects of solid fuel ramjet (SFRJ) are reviewed. On the point of view of the ability of an SFRJ to operate satisfactorily at all off-design conditions the areas of concern to propulsion system designer are (1) selection of a fuel type, (2) flame holding requirements that limit maximum fuel loading, (3) understanding the fuel regression rate behaviour as a function of flight speed and altitude, (4) diffusion-controlled combustion process and its efficiency enhancement, and (5) inlet/combustor matching. Considering these areas, the following aspects are reviewed from the information available in open literature: (1) different experimental set-up conditions adopted in combustor research, (2) various suitable fuel types, (3) flammability limits, (4) fuel regression rate behaviour, (5) methods of achieving high efficiency in metallized fuel, and (6) various modelling efforts. Detailed discussion is presented on two different types of regression rate mechanism in SFRJ: one that is controlled by the heat transfer processes downstream of the reattachment region and the other by that in the region itself. With a view to demonstrate the use of the information collected through this review, a preliminary design procedure is presented for an SFRJ-assisted gun launched projectile of pseudo-vacuum trajectory. © 1998. Published by Elsevier Science Ltd.

Journal	Journal of Propulsion and Power
ISSN	07484658
Open Access	No