Combustion Characteristics of a Paraffin-Based Fuel Hybrid Rocket Tsong-Sheng Lee 1 and Hsin-Luen Tsai 2*

Transcription

1 9 th Asia-Pacific International Symposium on Combustion and Energy Utilization Combustion Characteristics of a Paraffin-Based Fuel Hybrid Rocket Tsong-Sheng Lee and Hsin-Luen Tsai * Aerospace Science and Technology Research Center National Cheng Kung University, Tainan, Taiwan Department of Electronic Engineering Kao Yuan University, Kaohsiung County, Taiwan *corresponding author: hsinluen@gmail.com Abstract The hybrid rocket has the superiority of safety, low-cost and high-reliability over solid-propellant and liquid-propellant rockets. Classical hybrids, however have suffered from slow fuel regression rates and relatively poor combustion efficiency. Enhancement effects of fuel composition and swirling oxidizer injection on the fuel regression rates were investigated in the present experimental study. Various fuel grains with different weight ratios of paraffin wax and HTPB mixtures including 9P (9% paraffin + % HTPB), 7P and 5P were developed. Grain strength was dropped significantly as test motor was ignited for the 9P fuel. The 7P fuel could be maintained stable combustion under most test conditions except at higher oxygen flux test in which pressure fluctuations was observed, and part of fuel vapor was blow with exhaust flame. With swirling gaseous oxidizer (GO) injection, the 5P fuel has shown about 6% increase in fuel regression rate as compared to classical HTPB/GO hybrid rocket. Moreover, 5P fuel motor could also be operated at fuel rich combustion condition that would reduce graphic nozzle erosion effect. With the measured chamber pressure and oxygen mass flow rate and the calculated fuel characteristic velocity, histories of the spatially averaged of regression rate, oxygen mass flux, oxidizer/fuel ratio, and thrust coefficient were obtained by employing mass conservation law. Due to the variation of port diameter, wide range of oxygen mass flux and regression rate could be obtained during single test and the data reduction analysis has developed to be proven as an effective way to estimate fuel regression rate at single test. Internal ballistic analysis on combustion history inside a motor will probably provide as a guideline for aiding the design of a high performance hybrid rocket. These experimental findings indicate that paraffin-based fuel provides the opportunity to satisfy a broad range of mission requirements for the next generation of hybrid rockets. Keywords: hybrid rocket, paraffin-based fuel, fuel regression rate, internal ballistic analysis

2 Introduction A little bit lower performance as compare with solid rocket, hybrid rocket motor was however, finally chosen as the engine to fly SpaceShipOne to win the Ansari X Prize for the superiority of its low-cost and high-reliability. The success of SS could increase general interest in hybrid engines. Classical hybrid rockets have suffered from slow solid-fuel regression rates and relatively poor combustion efficiency. The complete modeling of hybrid motor combustion is quite complicated due to various physical and chemical processes. The model has to consider in a fuel grain passage a reacting flow created by the two distinctly different fluids: one, the mostly-vaporized- oxidizer entering the fore end of the fuel grain passage and the other, the fuel vapor blowing from the passage-wall. The boundary layer growing from the fore end of the passage contains the diffusion flame front within. Fuel is vaporized as a result of heat transferred from the flame front to the fuel surface. The fuel vapor converts towards the flame front while the oxidizer from the free stream diffuses into the boundary also towards the flame front from the opposite direction. Classical hybrid rockets have not yet found however, widespread use for either commercial or military applications, possibly because they suffer from slow fuel regression rates, low volumetric loading, and relatively poor combustion efficiency. To achieve the necessary mass flow rate of pyrolyzed vapor from the fuel grain to produce the desired thrust level, complex cross sectional geometries with large wetted surface are must be employed. Such grains require large cases and display poor volumetric efficiency. Low regression rates are also disadvantage when small L/D ratios may be desirable, such as for upper stages. The limit on regression rate for the conventional hybrid combustion configuration is set by the physical phenomena of heat and mass transfer from the relatively remote flame zone to the fuel surface. As a consequence, the regression rates of modern hybrids that utilize polymers as the fuel are much lower than conventional solid-rocket burning rates [-4]. Various methods for increasing fuel regression rates have been suggested in the past. The addition of ammonium perchlorate(ap)and/or aluminum in HTPB fuel is able to enhance the regression enhance rate at high GOx mass flux conditions []. A vortex hybrid engine having the capability of generating coaxial, co-swirling, counter flowing vortex combustion field designed by Knuth et.al [9]. Tremendous enhancement of fuel regression rates was obtained up to 65% larger than those in similar classical hybrids. By employing swirling GOx injection technique, Lee [8] demonstrated 5% regression rate and % Isp (specific impulse) increases as compared to non-swirling one in a HTPB/GOx hybrid motor system. Increased regression rates by several hundred percent have been observed with solid cryogenic hybrids including several frozen organic liquids and normal pentane. The regression-rate model for these liquefying fuels has been developed by Karabeyoglu [6]. Very high regression rates observed in the cryogenic tests have been successfully predicted by this liquid layer theory which also led to the conclusion that paraffin waxes will exhibit high regression rates comparable to pentane. In order to improve the grain strength during the heating period, the fuel is fabricated by mixing HTPB with the paraffin. A series of paraffin-based hybrid rocket fuels were tested in laboratory-scale motor that was designed similar to the classical one with an aft-mixing chamber. By applying the oxidizer swirling inflow to improve the mixing mechanisms between the oxidizer and fuel vapor, fuel regression rate and specific impulse were selected as the indicators for the performance of the developed hybrid rocket fuel. It was observed that there existed some difficulty for increasing specific impulse effectively in a hybrid rocket system, probably because of the complicated combustion behavior inside the motor. Internal ballistics of a hybrid rocket was therefore, also investigated to examine the factors influence the performance of a hybrid rocket. Experimental Setup Fuel grain specimen Paraffin wax and HTPB were selected as the solid fuel in the present study. Paraffin is an odorless, tasteless, waxy solid. It melts between 47 C and 65 C and is unaffected by most common chemical reagents but burns readily in air. Chemically, paraffin is a mixture of high molecular weight alkanes with the general formula, where n is an integer between and 7. HTPB (Hydroxyl-terminated Polybutadiene) cross-linked with isophorone diisocyanate, IPDI has widely been used as the solid fuel for hybrid rocket motors. Weight ratios of 5% paraffin and 5% HTPB fuel grain was prepared by mixing liquefied paraffin with HTPB/ IPDI(9/8) mixture at 8C. This mixture was then poured into cylindrical motor equipped with a mode for tailoring the grain configuration. With same curing process as carried out as for HTPB grain [8], this case bonding fuel grain, 4mm diameter*8mm length, is ready for test. The grain configuration, as shown in Fig., that consists of a mm diameter*mm length flow passage and a divergent perforation section of 7mm with half angle of.4 degree and mean diameter of 7.7mm was employed in the present study. The spe-

3 cific weight of the test grain is.9. Test Equipment The hybrid rocket test facility consists of Gaseous Oxygen supply, N purge, pyrogen ignition and thrust measuring systems. Oxygen is supplied from bank oxygen bottles kept at maximum pressure of 5 Mpa. A pneumatic control ball valve located on thrust stand is used to initiate and terminate the oxygen flow. A Venturi nozzle is placed in the line to measure mass flow rate. The pressure upstream the nozzle is always kept at a constant pressure through a regulator. Oxygen supply to the motor is through two hanged flexible Teflon hoses. Nitrogen is employed as purge gas to terminate combustion after the desired burning time. Propane gas and a spark formed the pyrogen igniter to initiate the test. Hybrid rocket thrust is measured by a flexure plate type thrust stand, in which hanged horses design and placed pneumatic control devices on the stand are employed. The natural frequency of the thrust stand measured is 8Hz. The operating pressure of the motor was controlled using interchangeable graphite exit nozzle. Oxygen was guided to enter the combustion through a swirler having tangential inlets on the circumference of the inner case and providing the swirling gaseous oxygen injection capability. Each inlet has diameter of mm hole and makes angle of 55 with motor axis. This annular swirler is then characterized as strong swirl one with swirl number of.95. Experimental Procedure The oxygen regulator located at upstream of the Venturi was first adjusted to supply the desired oxygen mass flow rate during test. The pyrogen igniter was initiated and maintained for about second. Immediately after turning off the ignition system, the pneumatic control valve was actuated to supply the oxygen. After the desired burning time, in quick successions oxygen supply was cut off and nitrogen purge was opened to extinguish combustion. The test measurements were the thrust and chamber pressure and pressure at upstream and the throat of Venturi. All of the signals from stain-gauge type, pressure transducers and oxygen temperature were recorded by LabView system. The sampling rate was samples/s. Other pre-and post- test measurements were the initial and final nozzle throat diameters and grain mass and dimension. For each test, the desired oxygen mass flow rate could be obtained by choosing an appropriate Venturi nozzle throat and its upstream pressure. Results and Discussions Hybrid rocket combustion can be characterized as diffusion flame that is mainly controlled by the degree of the mixing of oxidizer and vaporized fuel. The more chance provided for mixing, the better combustion performance of a hybrid motor would be. To enhance the mixing and flame holding capability, the idea of providing re-circulation zone in the taper fuel grain design is employed in the present study. The location of 5mm diameter just formed as a sudden expansion act as flame holder in the combustion regime Data Reduction The primary observed variables in the test were the time history of chamber pressure and thrust, fuel consumed and oxidizer mass flow rate which was held constant. With the measured values of the thrust, chamber and the oxygen mass flow rate, the spatially averaged instantaneous regression rate, port diameter were obtained by employing the conservation scheme from reference []. Specific impulse Isp was evaluated based on burning time, T b that was defined as the tangent bisector point on the pressure-time trace. Fuel regression rate was calculated based on space-time-averaged scheme given by: r = ( rf r ) / T b Where r and r f were port radius evaluated at time t = and t = T b The averaged oxidizer mass flux GOx is estimated based on the average port radius: 4mo GOx = π ( r + rf ) Combustion Performance of HTPB and 9P In order to improve the performance of a hybrid motor, various weight ratios of paraffin wax and HTPB mixture have been developed and tested. HTPB is the binder of solid propellant and also selected as the hybrid rocket fuel grain. At the higher oxidizer flow flux, the combustion is easy to lead to the lean burning due to the lower fuel regression rate, thus the excess oxidizer will combust with the graphite nozzle throat and cause the erosion. Fig. shows the jet plume from the nozzle exit and demonstrates the obvious shock trains along the downstream at the oxidizer GOx= kg/m -sec. After heating, the paraffin-based fuel becomes molten easily and its higher regression rate results in the fuel-rich combustion characteristics. Fig. shows the jet plume

4 ejected from the nozzle during the 9P fuel grain testing. For comparison with the literature results, under non-swirling oxidizer inlet conditions, the fuel regression rate test in Fig. shows the relation of HTPB fuel regression rate against the oxidizer supply is very consistent with the empirical laws derived from Sutton []. Fig. 4 shows the fuel regression rate of 9P paraffin-based fuel is slightly deviated, but in the same trend as the results from Stanford group [6]. Combustion Characteristics of 5P fuel According to the results from the experimental fuel testing, the 5P fuel with the mixture of 5% paraffin wax and 5% HTPB fuel could keep structural integrity of fuel grain after heating and fuel testing periods, and will meet the requirements of increasing the fuel regression rate and reduce the graphite erosion of rocket nozzle. Fuel regression which was assumed to be a constant providing a constant GOx supplying was estimated mainly by the weight method. Fig. 5 shows the regression rate and specific impulse (Isp) of 5P fuel under different oxidizer rates with swirling and non-swirling conditions. By observing the results, the fuel regression rate increases 65% and specific impulse increases 5% at GOx= kg/m -sec due to the swirling effect. The Isp could reach to sec at GOx=5 kg/m -sec under the swirling condition. These hybrid fuel tests demonstrate the swirling conditions will results in better mixing between fuel vapor and oxidizer, and combustible mixture, so improve the fuel regression rate and specific impulse. According to the experimental results of 5P fuel, under the swirling condition, the fuel regression rate expression (mm/sec) for 5P fuel grain is expressed as the power expression of GOx, where GOx is oxidizer flux in kg/m-sec,.876 r =.6 GOx Performance analysis of hybrid rocket The present study adopts the internal ballistic analysis developed by Philmon [] to calculate whether the total fuel mass conserved for every time step by assuming the combustion efficiency kept constant and investigate the internal ballistics characteristics during the hybrid rocket tests. The mixture of 5% paraffin wax and 5% HTPB fuel, 5P presented the best performance among all test fuels and then selected as the fuel for the following testing. Fig. 6 shows the static test results including thrust, upstream pressure, and chamber pressure etc. against the test time under the 5P hybrid fuel burning test with swirling oxidizer inlet. () () Fuel grain port area and equivalence ratio: Combustion of HTPB and paraffin wax (n=7) can be simply expressed as: H O 4CO + H O C4 6 C7H O 7CO + 8H The stochiometric mixtures of fuel and oxygen ratio for these two fuels are.7 and.9 respectively. The ratio of.98 is employed for the fuel of 5% paraffin wax and5% HTPB mixture. Instantaneous equivalence ratio at t i is, therefore defined as φ i = m f, i / m o, i /.98 With the measured chamber and the oxygen mass flow rate data and the calculated fuel characteristic velocity C* the instantaneous spatially averaged equivalence ratio could be estimated under the constant combustion efficiency assumption. The combustion efficiency is kept as.97 to meet the iteration requirement of internal ballistics analysis. Fig. 7 shows the change of the calculated fuel port area and equivalence ratio along the burning time. The test uses the propane gas and a spark to ignite the hybrid rocket, and ignition gas is supplied for.5 second for heating. As the gaseous oxidizer supplies, there are amount of high temperature fuel vapor accumulated near the fuel grain surface. So as to the early period of hybrid fuel combustion about.7 sec, the calculation is not converged because the program can find the adequate equivalence ratio for mass conservation. The consumed fuel mass within.7 second is estimated by weighting the ratio of the impulse in.7 second to total impulse, thus the initial port diameter is also estimated as.8 mm, not the original one.5 mm. By observing the equivalence ratio figure, the value is above unity, thus indicated the combustion is in fuel rich condition. Fig. 8 show the thrust coefficient is compared with the theoretical one where the erosion does not occur in this test. If the erosion occurs in the nozzle throat, the chamber pressure will decrease, thus the thrust coefficient increase accordingly. So in the present analysis, the nozzle throat variation is taken into account to keep the thrust coefficient as the theoretical one. () Fuel regression rate and oxidizer mass flux Fig. 9 shows the time histories of gaseous oxidizer mass flux and fuel regression rate. At early period of combustion, the oxidizer mass flux is higher due to the smaller fuel port diameter, and decreases as the combustion proceeds and the fuel grain depletion occurs. And the higher fuel regression rate occurs at the early period because of the higher oxidizer mass flux and high heat flux from the ignition combustion mixture. O

5 () Different oxidizer mass fluxes Fig. shows times histories of the equivalence ratio under different oxidizer mass fluxes by testing 5P hybrid fuel. The figure exhibits that at lower oxidizer mass flux GOx=89~7 the combustion belongs to be fuel rich and becomes towards fuel lean by increasing the oxidizer mass flux. As the oxidizer mass flux at GOx=4, the combustion is fuel-lean partially. (4) Different fuel compositions The time histories of equivalence ratio for two different fuel compositions, 5P and 7P is shown in Fig. for both swirling and non-swirling oxidizer injection techniques (GOx=5~45). This figure shows that fuel rich combustion occurred at early period of the testing for all fuel tests. This is probably caused by the molten liquid layer, which was formed by the pyrogen ignition high temperature gas, subjected to large shear force generated by high oxygen mass flux. For non-swirling cases, the combustion for both of fuel compositions are fuel-lean, and the combustion of 7P fuel is fuel-rich at the early period. By employing the swirling oxidizer injections, the combustion becomes fuel-rich for both cases and indicates the swirling oxidizer injection will enhance the fuel regression rate. Furthermore, the 7P fuel is more easily vaporized due to the higher paraffin was content and its equivalence ratio is higher than the cases of 5 P fuel. (5) Time histories of fuel regression rate Basically under constant oxidizer mass flux, the fuel port diameter increases as combustion proceeds, but eh fuel regression rate decreases as time evolves. By the present internal ballistics analysis, the fuel regression rate expression as function of oxidizer mass flux could be obtained by single experimental results without many experiments. Fig. shows the fuel regression rate comparison of equation () and two experimental results (GOx=89 and ) calculated by the present method and their results are consistent with the equation () very well. Fig. shows the internal ballistics analysis of 5P fuel for three different oxidizer mass flux and the results are also consistent with equation (). The present data reduction method is applicable to reduce the requirement of experimental burning tests and costs. Conclusions A series of paraffin-based hybrid rocket fuels have been studied experimentally in a laboratory-scale motor that was designed similar to the classical one with an aft-mixing chamber. To enhance the fuel and oxidizer mixing level inside the motor, technique of swirling GOx injection was employed. Internal ballistics of a hybrid rocket burning with different weight ratios of paraffin wax and HTPB fuel were investigated. Information of combustion behavior inside the motor will probably provide as a guideline for designing a high performance hybrid rocket. Several conclusions of the present study are listed as follows.. Hybrid fuel grain with different weight ratios of paraffin wax and HTPB are fabricated as P, 9P, 7, 5P and pure HTPB. In contrast to HTPB, high Paraffin contented fuel easily becomes molten and results in high regression rate. For 9P and 7P fuels, the structural strength of fuel grains are not strong enough to sustain and parts of the liquid molten fuels are ejected from the nozzle without reacting with the oxidizer. These issues limit their application as hybrid rocket fuel. For 5P fuel, it still keeps its structural integrity after heating and little erosion impact on the nozzle throat after fuel test. The 5P fuel regression rate expression against oxidizer mass flux is expressed.876 as r =.6 GOx and increase 6% as compared to the pure HTPB fuel under swirling condition.. By assuming the combustion efficiency kept constant and the measured chamber pressure and thrust, the data reduction method based on the mass conservation is developed to calculate the fuel regression rate, oxidizer mass flux, specific impulse, equivalence ratio, thrust coefficient etc. to investigate the complex combustion characteristics for designing the next generation propulsion systems. The present internal ballistics method is applicable to obtain the relation of fuel regression rate and oxidizer mass flux by single burning test.. The mixture of 5% paraffin wax and5% HTPB, 5P fuel presented the best performance among all test fuels. About 65% regression rates and 5% specific impulse increased by burning with 5P fuel, due to the swirling effect at GOx=. Specific impulse of 5P fuel can reach up to sec as GOx=5. Paraffin-based fuel provides the opportunity to satisfy a broad range of mission requirements for the next generation of hybrid rockets Acknowledgement The authors acknowledge the support provided by National Science Council, under the contracts NSC95--E-6-6 and D. The authors also thank the technical support of ASTRC/NCKU engineers with the experiments.

6 References [] [] [] [4] Heslouin, A., Simon, P., Lengell, G., Foucaud, R., Gibek, I. And Pillet, H., Propulsion of Microsattllites by Hybrid Rocket Engines, st International Hydrogen Peroxide Propulsion Conference, The University of Surrey, UK., July 9-, 988. [5] A Moore, G. E., and Berman, K., A Solid-Liquid Rocket Propellant System, Jet Propulsion, pp , Nov [6] Karabeyoglu, M. A., Altman, D., and Cantwell, B. J., Combustion of Liquefying Hybrid Propellants: Part, General Theory, Journal of Propulsion and Power, Vol. 8, No., pp6-6, May-June. [7] Yuasa, S., Shimada, O., Imamura, T., Tamura, T., and Yamamoto, K, A Technique for Improving the Performance of Hybrid Rocket Engines, AAIA paper 99-, 99. [8] Lee, T. S. and Potapkin, A., The Performance of a Hybrid Rocket with Swirling GOx Injection, Proceedings of International Conference on the Methods of Aerophysical Research, Part, Novosibirsk, Russia, pp.6-, July. [9] Knuth, W. H., Chiaverini, M. J., Sauer, A., and Gramer, D.J., Solid Fuel Regression Rate Behavior of Vortex Hybrid Rocket Engines, Journal of Propulsion and Power, Vol. 8, No., pp.6-69, May-June. [] Sutton, G.P., Rocket Propulsion Elements, 6th edition, ISBN , 99. [] Philmon,Georage et.al., Fuel Regression Rate in HTPB /Gaseous-Oxygen Hybrid Rocket Motors, pp5-4, J. of Propulsion and Power, Vol.7, No., Jan-Feb.. [] Lee,T.S., andcho, S.M. Cho, Internal Ballistics of a Hybrid Rocket Burning with Paraffin-Based Fuels, 6 AASRC/CCAS Joint Conference, Dec,6. Fig. Combustion plume of % HTPB fuel at nozzle exhaust Fig. Combustion plume of 9P fuel at nozzle exhaust Regression rate of HTPB fuel on no swirling condition Sutton[] H GOx (Kg/m sec) Fig. Fuel regression rate of HTPB under non-swirling oxidizer injection

7 4 Regression rate of 9P fuel on no swirling condition Variation of port diamter and phi with time (No: 9654).5 Karabeyoglu[6] 9P Port diameter (mm) 5 5 D PHI.5.5 Equivalence ratio (phi).5 Fig GOx (Kg/m sec) Comparison of 9P fuel under non-swirling oxidizer injection Combustion character of 5P fuel Time (sec) Fig. 7.8 Fuel port diameter and equivalence ratio against time Thrust coefficient (Test No:9654) (5P+5H,GOx=7,Swirling) r(ns) r(s) ISP(NS) ISP(S) 5 5 ISP (sec) Thrust coefficient (Cf) Cf Cf(theory) GOx (Kg/m sec) Fig. 5 Fuel regression rate (r) and specific impulse (ISP) of 5P fuel (NS:non-swirling, S:swirling) Pressure (Kg/cm ),Thrust(Kg) 5 5 Static Firing Test 9654 (5P, GOx=7, Swirling) Pc thrust PP PP Test time (sec) Fig. 6 Static test measurement of 5P fuel burning test under swirling oxidizer injection Upstream P (Kg/cm )G GOx (Kg/m sec) Time (sec) Fig. 8 Thrust coefficient against time Variation of Oxygen Flux and Regression rate with time (Test No:9654) GOx r Time (sec) Fig. 9 Oxidizer mass flux and fuel regression rate against time.5.5.5

8 Equivalence ratio comparison for 5P+5H fuel in swirling condition Regression rate comparsion (5P+5H,Swirling).5.5 Equivalence ratio (PHI).5 GOx=89 GOx=7 GOx= GOx=4.5.5 Equ. () GOx=89 GOx= GOx= Dimensionless time Fig. Comparison of equivalence ratio for different oxidizer mass flux against time GOx (Kg/m sec> Fig. 5P fuel regression rate under different oxidizer mass fluxes Effect of fuel composition on equivalence ratio (GOx 5~45) Equivalence ratio (PHI).5.5 5P (S) 7P (S) 5P (NS) 7P (NS) Dimensionless time Fig. Comparison of equivalence ratio for two different fuel compositions.5 Regression rate comparison using experimental data.5.5 Equ () GOx=89 GOx=.5 Fig GOx (Kg/m sec) Comparison of fuel regression rates by single test measurement