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The Development of the Hybrid Rocket for Bloodhound Ssc

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The Development of the Hybrid Rocket for Bloodhound Ssc Copyright © 2012, The Falcon Project Ltd and Fluid Gravity Engineering Ltd THE DEVELOPMENT OF THE HYBRID ROCKET FOR BLOODHOUND SSC Jubb, D. ([email protected]); Merrifield, J. ([email protected]) The Falcon Project Ltd, Fluid Gravity Engineering Ltd UNITED KINGDOM 1. A BSTRACT BLOODHOUND SSC® aims to inspire the next generation of scientists, engineers and mathematicians by setting a 1,000 mph land speed record. The vehicle is powered by an EJ200 jet engine and a hybrid rocket. The type of rocket and the oxidiser selection were important safety considerations in the early stages of the programme. Figure 1: BLOODHOUND SSC The hybrid uses HTP, decomposed by a silver catalyst and an advanced fuel grain based on HTPB The primary objective of the programme is (including a small quantity of solid oxidiser and to address the shortage of engineers and scientists, aluminium powder). The 18-inch (457mm) diameter by producing an iconic project to inspire the next chamber has been developed using a 6-inch generation to consider careers in these areas. (152mm) diameter prototype, supported by Richard Noble OBE is the project director and Wg computational modelling. The hybrid chamber is the Cdr Andy Green OBE (the current land speed largest to be developed in the UK, producing an record holder) is the driver. average thrust of 111 kN for 20 seconds and a peak thrust of 122 kN. HTP is delivered to the chamber After an extensive evaluation of solid and using an up-rated version of the Stentor large HTP liquid propellant systems, a hybrid rocket was pump from the Blue Steel cruise missile. The pump selected as the preferred candidate. The next stage is driven by a Cosworth V8 Formula 1 engine, was propellant selection, Liquid Oxygen, Nitrous which also serves as the vehicle APU. To date, the Oxide and Nitric Acid were all considered, before development programme has included 11 static tests final selection of 87% High Test Peroxide (HTP). of the 6-inch (152mm) hybrid chamber, 3 static tests A HTPB based fuel grain was chosen after a of the 44.5 kN monopropellant chamber and a selection process which included PBAN and PVC. single static test of the 18-inch (457mm) hybrid chamber. High fuel grain regression rates have been 3. 6-INCH (152MM) C HAMBER achieved, along with uniform regression and exceptional combustion stability. Criteria have also Prior to development of the full size been established for the safety and acceptance chamber a 6-inch (152mm) diameter research testing of the rocket system for use in chamber was produced. A modest specific impulse BLOODHOUND SSC. target of 200 sec was set early in the programme to ensure that the vehicle performance estimates were 2. I NTRODUCTION based on conservative rocket performance. BLOODHOUND SSC is powered by an The Mk1 version has an aluminium EJ200 jet engine from Eurofighter Typhoon and the combustion chamber. Eleven static tests of the 6- largest hybrid rocket to have been developed in the inch (152mm) chamber have been conducted to UK. The technical challenge of designing a vehicle date. The objectives were to refine the: which can achieve 1,000 mph (1,610 kph) on land is very substantial, in terms of both aerodynamic • Fuel grain geometry design and propulsion. These challenges are being used to highlight key aspects of the curriculum • Fuel grain composition through an education programme, which now includes over 5,000 schools, colleges and • Catalyst pack, and universities in the UK. • Thermal insulation The Development of the Hybrid Rocket for BLOODHOUND SSC Page 1 of 6 Copyright © 2012, The Falcon Project Ltd and Fluid Gravity Engineering Ltd The catalyst pack is located at the top of the combustion chamber and consists of 80 silver plated nickel gauzes, similar to the type used in the Gamma series of rocket engines used in the British Black Arrow launch vehicle. The general arrangement of the 6-inch (152mm) chamber is shown in Figure 2. The 6-inch (152mm) chamber has proved to be an excellent development tool. Early tests focused on the length of the fuel grain and post fuel grain mixing chamber. A conventional 9-point star Figure 3: Firing 010 of the 6-inch (152mm) configuration was selected for the fuel grain. The Hybrid Chamber use of a small percentage of ammonium perchlorate and various burning rate modifiers in the fuel grain, 4. C OMPUTATIONAL M ODEL combined with careful control over the fuel grain density, allows high regression rates to be achieved. Fluid Gravity Engineering’s TINA code was The fuel grain composition was also tailored to used to provide CFD support to the hybrid rocket provide a rough surface during the firing, this development for BLOODHOUND SSC. One of the promotes local turbulence improving mixing and main difficulties concerned with predicting hybrid entrainment. Two of the static tests (006 and 008) rocket performance is the highly coupled interplay resulted in chamber failures. Both failures were between oxidiser mass flow rate and fuel grain associated with the thermal insulation around the vapour liberation. The close coupling between catalyst pack/fuel grain interface. Additional silica viscous, inertial and chemical effects makes the phenolic insulation was added in this area and this problem of hybrid rocket performance a good performed well in subsequent firings. candidate for application to CFD, since this interaction can lead to large errors in simplified The Mk2 chamber is currently under treatments of hybrid rocket performance. development, it has a carbon fibre combustion chamber, an inconel catalyst pack body/injector and The key code development was to implement a lightweight nozzle. a boundary condition appropriate to the hybrid rocket motor problem. In particular the boundary condition is designed to release the correct amount of fuel vapour from the fuel grain surface into the flow, based on heat flux to the surface. In order to tackle this problem, two main fluid properties need to be predicted at the fuel grain surface. These are temperature and gas composition. These properties are calculated using an energy and mass balance relationship. The temperature at the surface is used to uniquely define the rate at which fuel vapour is liberated through the empirical relationship shown in Equation.(1). Figure 2: Sectioned Mk2 6-inch (152mm) Hybrid Chamber ⎛⎞−Ea (1) ρsbrA= sexp⎜⎟ ⎝⎠RTu where rb is the fuel grain regression rate, ρs is the fuel grain density, Ru is the universal gas constant, T is the surface pressure and Ea and As are empirical constants. These constants can be obtained for HTPB from the works of Brill[i] or Cohen[ii]. The Development of the Hybrid Rocket for BLOODHOUND SSC Page 2 of 6 Copyright © 2012, The Falcon Project Ltd and Fluid Gravity Engineering Ltd The energy balance which must be solved can This treatment of the oxidiser/fuel grain be written as surface interaction was used to assess rocket motor performance to support the 6-inch (152mm) ∂T prototype configuration. In this example, we were (2) −=κ Q ∂y cond required to simulate performance near the start of the fuel grain burn. Computational resources and time constraints were such, that only two where κ is the thermal conductivity of the gas dimensional simulations were possible. In order to at the surface of the fuel grain, y is the direction address this problem, an equivalent two dimensional perpendicular to the grain surface, and Qcond is the geometry relevant to the start of the fuel grain burn heat flux being conducted away from the surface was used. The equivalent geometry was chosen so into the fuel grain. The conduction term can be that the following physical properties were matched calculated by an external FEM or if the fuel grain is with the real geometry: assumed to be thick and heat fluxes are not varying too quickly in time, a steady state assumption can • Exposed fuel grain area (start of burn) be used to simplify the calculation of Qcond such that • Fuel grain inflow area (3) QrHHcond=−ρ s b() T 300 • Nozzle expansion ratio where H is the enthalpy of the solid fuel • Nozzle exit plane area grain. The subscript T represents evaluation at the surface temperature while the subscript 300 In order to match these properties in 2D, an represents evaluation at the ambient temperature annular combustion chamber and nozzle were (300K). The full simplified energy balance can then modelled. The inner wall of this configuration is be written as inviscid (i.e. does not contribute to thermal and viscous losses). Since the mixing chamber had been identified as an important contributor to motor ∂T ⎛⎞−Ea (4) −=κ AHHsTexp⎜⎟() −300 . performance, a mixing chamber was added before ∂yRT⎝⎠u the nozzle. The outline for this representative geometry is given in Figure 4. This is the form of the energy balance equation implemented within TINA in support of the BLOODHOUND SSC project. The mass balance that must be solved can be written as ⎛⎞−Ea ∂Y (5) ρsrX b pyrol==−+ A sexp⎜⎟ X pyrolρρ D v g X ⎝⎠RYu ∂ y where Xpyrol are the pyrolysis gas species mass fraction, Y are the gas phase species mole fractions, D are the gas phase diffusion coefficients, X are the gas phase species mass fractions and vg is the gassing velocity. The gassing velocity is calculated using mass conservation such that (6) ρvg = ρ s rb . Figure 4: Representative 2D geometry Equations (4) and (5) are iterated using a Details of the temperature map obtained Newton Raphson scheme to solve for T and Y (all through the nozzle and mixing chamber are given in other surface variables can be calculated from Figure 5.
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