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Understanding the Rocket Engine Performance in BLOODHOUND SSC the BLOODHOUND Engineering Project Mechanical Engineering

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Understanding the Rocket Engine Performance in BLOODHOUND SSC The BLOODHOUND Engineering Project Mechanical Engineering This exemplar requires some knowledge of Physics and Chemistry together with Maths. INTRODUCTION SCENARIO Figure 1: The BLOODHOUND SSC The BLOODHOUND Project is an iconic engineering and education adventure that is pushing technology to its limit, providing us with a once in a lifetime opportunity to inspire the next generation of scientists and engineers. BLOODHOUND SSC (Super Sonic Car) has been designed to set a new land speed record of 1,000 miles per hour (mph). BLOODHOUND SSC is powered by a EUROJET EJ200 jet engine and a hybrid rocket engine. A rocket is basically an engine which carries both a fuel and oxidiser (it does not require any Figure 2: Static tests of the 15.2cm (6-inch) hybrid oxygen from the air). Hybrid rockets use a solid chamber fuel and a liquid oxidiser. The fuel is contained within the combustion chamber and the liquid The two pictures above in figure 2 show static oxidiser is injected at the top of the chamber. tests of the 15.2cm (6-inch) diameter hybrid Therefore a hybrid can be easily shutdown by rocket chamber that is being used to develop the turning off the supply of liquid oxidiser, making hybrid rocket for BLOODHOUND SSC. them more suitable for use in a land speed The hybrid rocket in BLOODHOUND SSC uses record car. Hybrids use the same oxidisers as HTP (a concentrated hydrogen peroxide H2O2, liquid propellant systems; fuels can include basically water with an extra oxygen atom) as synthetic rubbers and plastics (such as PVC). the oxidiser and a synthetic rubber Hydroxyl- IMPORTANCE OF EXEMPLAR IN REAL LIFE Terminated Polybutadiene (HTPB) as the primary fuel. The hybrid combustion chamber for Thrust is a reaction force described BLOODHOUND SSC has a catalyst pack to quantitatively by Newton's Second and Third decompose the HTP, the decomposition Laws. When a system expels or accelerates products then enter the fuel grain. The fuel mass in one direction, the accelerated mass will automatically ignites in contact with the exert an equal but opposite force on that system. decomposition products, generating pressure. A rocket engine produces thrust by expelling hot The combustion products then enter the nozzle gas at a very high velocity. which converts the high pressure, low velocity Calculating thrust and specific impulse are gas into high velocity low pressure gas fundamental to the design of any rocket engine. generating up to 122 kN (27,500 lbs) of thrust. Achieving the optimum oxidiser:fuel (O:F) ratio The chamber contains 181 kg of fuel which can and understanding how the chamber pressure run for up to 20 seconds. The chamber is 45.7cm and the combustion process affect specific in diameter and 4.27m long. Development work impulse is the key to achieving an efficient has been conducted on a 15.2 cm (6-inch) system. diameter chamber. Several firings have been conducted to test various aspects of chamber 1 design, fuel grain configuration and catalyst The second expression in equation (2) shows material. that the thrust produced by a rocket varies with altitude because of the dependence on ambient One of the challenges is to find optimum specific pressure. Thus, rocket performance is usually impulse of the rocket while reducing the defined by two limits T (thrust at sea level) and harshness of the environment (in terms of high sl heat fluxes) experienced by key components of Tv (thrust in vacuum). The customary index of the motor (specifically the nozzle and throat). performance is specific impulse, Isp defined by: Calculating the heat transfer rate to rocket motor T components is frequently dealt with using CFD I = … (3) analysis and is not discussed here. Also, there sp dm g × are so many dynamic variables in a rocket dt combustion chamber that we need to simplify the where g is the acceleration due to gravity which equations in order to get an approximate mathematical answer. To try and calculate all the varies with altitude in the same way that T does. dynamic variables requires extremely complex Using specific impulse to characterise the rocket mathematics. performance is analogous to using miles per gallon or specific horsepower to characterise a Therefore, in this exemplar, we will look at some car. of the fundamental aspects of calculating rocket performance and examine some of the more In order to address the question of O:F, we need complex aspects which are key to achieving an to keep some other parameters constant. The efficient hybrid rocket for BLOODHOUND SSC. mathematical analysis is simplest if we assume that the motor designers expand the rocket MATHEMATICAL MODEL engine nozzle in such a way that the static A rocket is propelled forward by a thrust force pressure at exit ( Pe ) closely matches the static equal in magnitude, but opposite in direction, to pressure at the operating altitude ( P ). Hence, the time-rate of momentum change of the a exhaust gas accelerated from the combustion the second term in equation (2) is usually small. chamber through the rocket engine nozzle Thus, ignoring that term, we get: (shown in figure 2). dm T = v … (4) dt e Mathematically, it can be expressed as: Combining equation (3) and (4), we get: dm v T = v … (1) I = e or I ∝ v … (5) dt sp g sp e where This shows that Isp is strongly dependent on the T = Thrust exit velocity ve which can be calculated from the dm = Mass flow rate of exhaust following expression: dt ⎡ ()γ −1 ⎤ Speed of the exhaust gases measured 2γ ⎛ P ⎞ γ v = v RT ⎢1 ⎜ e ⎟ ⎥ relative to the rocket e = c ⎢ − ⎜ ⎟ ⎥ … (6) ()γ −1 ⎝ Pc ⎠ Please note that equation (1) is true only for a ⎣⎢ ⎦⎥ perfectly expanded nozzle. where In case of BLOODHOUND SSC, this thrust is γ = Ratio of specific heats expressed as follows: R = Specific gas constant dm T = v + ()P − P A … (2) T = Combustion chamber temperature dt e e a e c where Pc = Combustion chamber pressure (Please refer to the isentropic expansion explained in ve = Exit velocity of the exhaust gases Reference 3, 4 or 5 under the title “Where to find more” in order to get better understanding of equation (6) above. Pe = Rocket engine nozzle exit pressure Proof of this equation is beyond the scope of this exemplar.) Pa = Ambient pressure For a particular fuel:oxidiser combination, γ is a A = Exit plane area slowly varying function of O:F and hence can be e considered as constant for simplicity. The ratio Pe Pc can be considered as a constant here to 2 get a fair approximation of the specific impulse. combustibles are here by design to enhance Isp , However, variable ratio P P requires more e c i.e. they should not be regarded as unburnt fuel. complicated analysis and is not covered here. Isp is the main driver of hybrid rocket design Thus, from equation (6), we see that: rather than combustion efficiency. Hybrid rockets are often run oxidiser rich (O:F larger than ve ∝ RTc … (7) stoichiometric) since a large flow rate of oxidiser We know that the specific gas constant R is is required to liberate fuel vapour from the fuel defined as the ratio of the universal gas constant grain. As a consequence, the Isp from a hybrid is R and the mean molecular weight , i.e.: u Mm usually lower than a high efficiency liquid engine R (cryogenic H2/LOX). R = u Mm CONCLUSION Hence, In this exemplar, we have looked at fundamental rocket performance parameters such as 1 R ∝ … (8) calculating thrust and specific impulse and also Mm made an attempt to explain why rocket exhaust plumes frequently afterburn upon mixing with the where Mm is the mean molecular weight of the ambient air producing bright flames in the combustion products in this case. exhaust gas. This is because optimum specific Combining equation (7) and (8), we can say that: impulse is frequently achieved when light flammable elements remain present in the Tc exhaust plume i.e. not all the fuel is burnt down ve ∝ … (9) to its final heavier state (usually H2O or CO2). Mm WHERE TO FIND MORE Finally, observing equation (5) and (9) together, it is clear that the optimum I can only be 1. Basic Engineering Mathematics, John Bird, sp 2007, published by Elsevier Ltd. achieved when the ratio of the combustion 2. Engineering Mathematics, Fifth Edition, John temperature Tc to the mean molecular weight Bird, 2007, published by Elsevier Ltd. Mm of the combustion products is maximised. In 3. Mechanics of Fluids, B. S. Massey, need particular, this reasoning explains why many more details. rockets produce afterburning plumes. At first 4. Rocket Exhaust Plume Phenomenology, glance, the combustion of H2 or CO in the plume Frederick S. Simmons, 2000, Published by would appear to point the inefficiency of the the AIAA rocket engine since burning these materials within the rocket engine would result in the 5. Rocket Propulsion Elements, George P. release of more chemical energy. However, the Sutton , Donald M. Ross, Oscar Biblarz; 7th mean molecular weight of a hydrogen oxygen Edition mixture is considerably lower than for pure water 6. http://www.bloodhoundssc.com vapour. As such, we find that an optimum Isp is 7. http://moodle.student.cnwl.ac.uk/moodledata achieved at a value of O:F (oxidiser to fuel ratio) _shared/CDX%20eTextbook/dswmedia/fuelS that is significantly lower than the stoichiometric ys/gasoline/fund/stoichiometricratio.html ratio (please refer to Reference 7) for complete combustion. This leads to the presence of combustible products in the plume.
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