[american institute of aeronautics and astronautics 41st aiaa/asme/sae/asee joint propulsion...

American Institute of Aeronautics and Astronautics

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90% Hydrogen Peroxide/Polyethylene Solid Fuel Hybrid Rocket Engine

Nobuo TSUJIKADO* Dept. of Aeronautics and Astronautics, Faculty of Engineering, Tokai University

1117 Kitakaname, Hiratsuka, Kanagawa 259-1292 JAPAN

Masatoshi KOSHIMAE, Rikiya ISHIKAWA and Kazuki KITAHARA† Gas Turbine Div., Kawasaki Heavy Industries, Ltd.

1-1 Kawasaki-cho, Akashi, Hyogo 673-8666 JAPAN

and

Atsushi ISHIHARA‡ Dept. of Mechanical Engineering, Saitama Institute of Technology

1690 Fusaiji, Okabe, Saitama 369-0293 JAPAN

The 90% hydrogen peroxide (HP), concentrated in lab-scale from Japanese domestic production of commercial grade HP, currently regulated below 60% for storing and handling safety have been applied to single perforation polyethylene solid fuel experimental hybrid rocket. Modified 3-way catalyst have been utilized for decomposing the HP. The temperature of the decomposed products of HP attain higher than auto-ignition with almost fuels, though a liquid rocket engine utilized for assisting solid fuel ignition and decreasing combustion pressure built-up time. In order to find oxidizer mass flux for sustaining hybrid rocket combustion, large fuel length/burning port diameter ratio (L/D=35) transparent polymethylmetacrylate solid fuel small hybrid rocket studies also have been carried out. The oxidizer mass flux have to be over the flux that extend diffusion flame up to the aft end of the solid fuel. The required minimum oxidizer flux were confirmed for the multi perforation solid fuel grain.

Nomenclature a = solid fuel burning surface regression rate constant in rb= a Gox

n [m/s]/[kg/m2-s] n Ab = solid fuel burning port surface area [m2] Ac = solid fuel burning port cross section [m2] Dp = solid fuel burning port diameter [m] Gox = oxidizer mass flux [kg/m2-s] Isp = specific impulse [N-s/kg] or [s] LDF = diffusion flame length formed on solid fuel burning surface [m] LSF = solid fuel length [m] mf = fuel mass flow rate [kg/s] mox = oxidizer mass flow rate [kg/s] mp = propellant mass flow rate: mp = mf + mox [kg/s] n = mass flux exponent of rb ; rb = a Gox

n rb = average solid fuel burning surface regression rate [m/s] ρf = solid fuel density [kg/m3] GOX = gaseous oxygen HP = hydrogen peroxide

* Lecturer (part time), Department of aeronautics and astronautics, [email protected] † Research Engineer, Division of Aeronautical Gas Turbine, [email protected] ‡ Associate Professor, Department of Mechanical Engineering, [email protected]

41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit10 - 13 July 2005, Tucson, Arizona

AIAA 2005-4091

Copyright © 2005 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.


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HTPB = hydroxyl terminated poly-butadiene LOX = liquid oxygen PE = polyethylene PMMA = polymethylmetacryrate

I. Introduction HE research and development works of hybrid rocket, which have been suspended these 30 years have been revived using mainly LOX or GOX and HP as oxidizer with HTPB or PE solid fuel for improving

environmental impact due to the rocket exhausts. The 80-90% HP had been applied to rocket engine of “Shusui” interceptor in Japan at the end of World War II. As liquid rocket oxidizer, comparing with 100% oxygen LOX, obtainable Isp are about 10% less than LOX as 90% HP contains only 29% oxygen though, the HP has several advantages such as that; 1. storable, not cryogenic is easy to handling, 2. about 22% higher density result in higher tank volumetric efficiency, 3, the optimum O/F ratio 7.2 is much higher than that of LOX oxidizer rocket. Therefore, in conjunction with the much lower than LOX rocket combustion temperature. Downsizing of combustion pressure which result in higher Isp, 4, extra low environmental impact combustion and exhaust products, etc. Among these advantages, combustion products contain a lot of low molecular mass H2O which brings not only relatively high Isp even at low combustion temperature as Isp ~ (Tc/M) (1/2), but also low environmental impacts, as shown in table 1.

In Japan, once produced 90% rocket grade (R/G) HP for rocket use in early 1960th but the productions have been discontinued except commercial grade (C/G) HP because of serious accident happened in concentrator. Now, the Japanese domestic productions of extra high purity and stable C/G HP for chemical industries are more than 300,000 tons/year (reduced 100%), its concentration still currently regulated below 60% for handing, storages and shipping safety, (The industries are planning to increase the concentration up to 70% in near future for saving about 15% ca. shipping cost.)

As the temperature of decomposition products of 70% (C/G) HP attains only some hundred degree C and contains about 10% liquid H2O even at high pressure, 70 % (C/G) HP oxidizer/PE solid fuel hybrid rocket have to be complicated special solid fuel igniter device1,2), though the obtainable Isp are about 240s, comparable to current high performance solid rocket propellant, In order to obtain 270s or more Isp, the (C/G) HP have been concentrated in laboratory scale for present 90 % HP/PE hybrid rocket studies. The firing tests have been conducted with experimental hybrid rocket engine, designed modeling current Japanese solid rocket booster, capable to test 1/3 (PE), 1/7.5 (PMMA) scale model single perforation solid fuel element and 1/10 scale model multi 7 perforations PE solid fuels. Using 1/7.5 scale model engine with transparent PMMA fuel, the combustion of the solid fuel burning port surface were observed. It was found that the HP oxidizer flow rate have to be sutisfied the oxidizer mass flux; Gox ≥ [4.a.ρf.(O/F)st .(L/D)][1/(1-n)] for sustaining stable diffusion flame formed on the solid fuel burning port inner surface.

Usually, the HP is decomposed passing through catalyst bed in advance to introduce into the solid fueled hybrid rocket combustion chamber. The temperature of the C/G HP decomposition products attains only 300ºC., so the solid fuel ignition device is indispensable1,2). However, in the case of R/G HP, the temperature exceed higher than 700ºC, the most of fuel will be ignited automatically, As the decomposition catalyst, piled up 70 to 130 silver net stack are widely utilized though, modified Pt. Pd 3-way catalyst are compact and active even for C/G HP.

II. Experimental Engine and Solid Fuel Design In order to design experimental hybrid rocket engine, a hybrid rocket booster and solid fuel grain were

temporarily designed modeling currently utilizing solid rocket boosters SRB-A for H-IIA and M-14 for M-V launch

T

Table 1. Characteristics of LOX, HP/PE Hybrid Rocket.3) Oxidizer (Density [kg/m3])

LOX (1,142.)

90% HP (1,395.)

70% HP (1,290.)

O2 Content [%] 100.0 29.24 21.65 (O/F)opt 2.53 7.2 10.4 Tc[K](@Pc[MPa]) 3,655.9 ( 7) 2,799.9 (10) 2,232.2 (10) Isp [N-s/kg] 2,647.7 2,689.4 2,232.2 (Isp [s]) (300.6) (274.2) (243.4) C* [m/s] 1,803.0 1,613.3 1446.8

Mol Fraction [%] of Combustion Products @ Combustion Chamber / Nozzle Exit CO 32.5/28.5 2.3/ 0.0 0.2/ 0.0CO2 14.1/21.0 15.4/17.6 13.2/13.0 H2O 33.3/39.8 78.0/81.1 86.9/86.5


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vehicle. The hybrid rocket booster HBRB-X, and booster rocket comparison are shown Fig. 1 and Table 2. The multi perforation type solid fuel grain for the HBRB-X also designed under the consideration of very low average burning surface regression rate compared to solid propellant about 0.33 [mm/s]. The cross section and fuel element are shown in Fig. 2. The fuel element is O.D.130/I.D.90, and 3,150 long PE cylinder.

In order to study the combustion characteristics of hybrid rocket, three type experimental engine were provided for firing tests (Fig. 3). The head block consisted of HP injector assembly, decomposition chamber and pre-combustion chamber were common to the engine.

The type I engine is for testing 1/7.5 scale model PMMA solid fuel element and aft combustion chamber for observing the combustion of fuel element. The transparent solid fuel served both as fuel and motor case employ the direct observation of burning surface. The type II engine is for testing 1/3 scale model solid fuel element. The type III engine is for testing 1/10 scale model 7 perforations solid fuel. The experimental solid fuel provided for the test were tabled below. (Table 3.)

The 3-way catalyst, designed for motor-bike was charged in the decomposition chamber. For C/G grade HP, original catalyst cartridge;38φ×150ℓ employed though, it was too much active and often dangerous for 90% HP. Therefore, decomposition test was conducted with about a half long cut of original cartridge (60/90ℓ) or the combination of 10ℓ and 25ℓ cut of the catalyst cartridge as shown left 2 catalysts arrangement in Fig. 4. The HP decomposition took place quickly and smoothly without liquid phase H2O at the test of below the HP flow rate 0.2 [kg/s]. However, the temperature of decomposed products did not attain to calculated 750 oC. due to the incomplete decomposition and also auto ignition of solid fuel did not take place. After all, solid fuel ignition are made with assistance of about 1 second duration liquid rocket igniter torch.

As long as utilize the igniter torch for igniting the solid fuel, the HP are not necessarily complete decomposition. Present decomposition chamber were employed shorter 19ℓ catalyst shown in Fig. 4. right.

Figure 1, Temporarily designed hybrid rocket booster.

Figure 2. Fuel grain cross section and fuel element.

Table 2. Booster Rocket Comparison Booster Rocket HBRB-X SRB-A M-14 Overall Diameter. [mm] 2,500 Overall Length [mm] 15,000 15,172 13,655Motor Gross Mass [kg] 78,000 75,000 80,000Propellant Mass [kg] 62,000 65,000 70,000Comb. Pressure [kPa] 10,000 11,760 4,900Avrage Thrust [kN] 24,500 20,100 38,700Effective Burn Time [s] 75 100 50Isp vac [N-s/kg] 2,990 2.745 2,705(Isp vac [s]) (305) (280) (276)

Figure 4. Decomposition catalyst arrangements set indecomposition chamber. (HP injected top center to down)

Table 3. Experimental Solid Fuel Dimension Length [mm] (--) not tested Fuel

Element Perfora

-tion L/D=35 L/D=30 L/D=20Full scale 90φ×1 (3,150) (2,700) (1,800)1/3 30φ×1 1,050 900 600 1/7.5 12φ×1 420 (360) (240)1/10 9φ×7 315 270 180


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III Firing Tests and Results The typical solid fuel

combustion in hybrid rocket are taken place from forward end and luminous diffusion flame last to the point where the oxygen in the oxidizer consumed as shown in Fig. 5. Since the diffusion flame are sustained with fuel generated on the burning surface by heat transfer from the flame, the diffusion flame have to be last up to the aft end of solid fuel or LDF ≥ LSF .

The average solid fuel burning surface regression rate is considered proportional to heat transfer rate and expressed as

rb = a Gox

n -----------------(1) a, n are estimated a=7.5±2.5.10-6 for PMMA, a=10±2.5.10-6 for PE, n=0.65~0.7 (calculated as 2/3) for both fuel, respectively. Fuel and oxidizer mass flow rate are expressed as ;

mf =π.Dp.LSF.ρf .rb= Dp.LSF ρf a Gox

n.-----------(2) mox= (π/4) .Dp2. Gox

-------------------------------(3) then the mixture ratio expressed as

O/F= mox/mf =[1/4 aρf ][Gox(1-n) /( LSF/Dp)]----(4)

Since the diffusion flame formed just under the boundary layer sustained nearly stoichiometric; (O/F)st combustion, the flame length is

LDF/Dp=[1/4 aρf ][Gox

(1-n) /(O/F)st]--------------(5) therefore oxidizer mass flux satisfy LDF ≥LSF is

Gox ≥ [1/4 a ρf (O/F)st (LSF/Dp)] [1/(1-n)]---------(6)

Substituting estimated n=2/3, necessary oxidizer mass flux Gox are rapidly increasing proportional LSF/Dp to the power third.

Two typical burning diffusion flames in the solid fuel burning port obtained Type I experimental engine firing test were showed in Fig. 6. The PMMA solid fuel is φ12/φ20×420ℓ; LSF/Dp=35 and calculated Gox lower limit nearly equal 400 [kg/m2-s]. Upper flame, the case mox=0.045[kg/s] (Gox≈395 [kg/m2-s]) was slightly below the lower limit though, the flame barely extended up to aft end of solid fuel and luminous plume was observed at nozzle exit. Contrary to upper photo., lower case, mox=0.035 [kg/s]; Gox≈395 [kg/m2-s] was far below the limit. Only sooty smoke without luminous plume was observed at nozzle exit.

Figure 5. Solid fuel combustion model.

Figur 3. Experimental hybrid rocket engines. ( A;Type-I, B -II and C;-III )

Figure 6. Typical solid fuel combustion. Oxidizer mass flux Gox≈395 (above), and Gox≈310 (bottom)f


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The firing test of Type II engine were conducted with 1/3 scale model single perforation experimental PE solid fuels; φ30/φ50×600ℓ, 900ℓ and 1,050ℓ (initial LSF/Dp=20, 30 and 35) and similar results were obtained. Fig. 7. shows typical firing test of 600ℓ solid fuel (initial LSF/Dp=20), mox=0.140 [kg/s] (Gox≈200 [kg/m2-s]) and P-T curve. Hard start pressure spike appeared in P-T curve though, as oxidizer mass flux greater than lower limit Gox=170 the hybrid rocket combustion were normally sustained. The solid fuel burning port diameter is increased

along with the progress of combustion. So, the oxidizer mass flux and burning surface regression rate are decreased though, since the burning surface area also increase, then fuel mass flow rate mf, total propellant flow rate mp and also combustion pressure are maintained almost constant level as shown Fig. 8. However, local regression rate along solid fuel were increased at right after down stream of forward end as seen in Fig. 9. due to the high heat transfer rate caused by steep temperature gradient formed by thin boundary layer. Therefore, the combustion pressure right after the ignition was slightly higher than successive pressure as seen in P-T curve of Fig. 7. Type III engine was provided for firing test of 1/10 scale model multi 7 perforations solid fuels. Typical firing test and P-T curve conducted with 180ℓ solid fuel were shown in Fig. 9. Different from single perforation solid fuel in Type II engine, since there was no combustion restriction wall the forward end surface of solid fuel was also burn same as port surface in Type III engine. Therefore, in order to achieve uniform ignition on the whole burning surface, pre-combustion chamber was enlarged and ignition plug also slightly sifted to down

Figure 9. Firing test of Type III engine with P-T curve.

Figure 7. Firing test of Type II engine with P-T curve.

Figure 8. Calculated parameter vs diameter. Figure 9. Local burning surface regression sketch.


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stream ward in Type III engine. These modification disappeared hard start ignition pressure spike as seen in Type II engine and obtained P-T curve ran smoothly as burning surface were widely distributed. Figure 10. shows forward and aft end surface after about 7 seconds firing test. The ignition and combustion of the forward end surface were normally carried out. The diameters of the forward end perforation were slightly larger than that of aft end similar as the residue of single perforation solid fuel of Type II engine test. Figure 11. is a roughly sketched local burning surface regression of the multi 7 perforation solid fuel.

The thrust of hybrid rocket is throttleable though oxidizer flow rate throttling. This is a strong point of the hybrid rocket system though, the throttling cause the diffusion flame to shrink and aft ward fuel combustion to run short, since the diffusion flame start from forward end and last up to the point that oxygen consumed under the nearly stoichiometric combustion. Figure 12. show residual solid fuel burnt for total about 40 seconds under the throttled oxidizer with original fuel and motor case. As seen in the figure, aft ward fuel burning were run short despite the forward completely burnt.

IV Conclusions Utilizing 90% HP as an oxidizer, concentrated domestic 60% commercial grade with one liter laboratory scale rotary evaporator, polyethylene solid fuel hybrid rocket including liquid rocket igniter torch have been evaluated. The yield of the single stage concentration was only attained 33% though, the yield was increasable 50% by 2 stages, 64% by 3 stages distillation. Since the yield of ideal concentration, bled off 1/3 H2O, is 67%, the yield of 3 stages concentration was not so low yield. The storing and handling stability of concentrated HP was kept at nearly same level of original 60% HP. The decomposition chamber and catalyst are key component for the HP rocket. Appropriately cut of 3-way catalyst cartridge designed for cleaning motor-bike engine exhaust have been applied for the present work, The decomposition were performed safely and much actively than usual silver net catalyst stack. In order to perform hybrid rocket solid fuel combustion, the length of diffusion flame formed on the solid fuel burning surface LDF has to be nearly equal or longer than solid fuel length LSF. For satisfying the requirement, oxidizer mass flow rate or mass flux has to be Gox ≥ [1/4 a ρf (O/F)st (LSF/Dp)] [1/(1-n)]. The thrust of the hybrid rocket is throttleable by the oxidizer throttling, though the throttling is undesirable because of the throttling cause aft ward solid fuel combustion to run short as the result of the decrease of mox, Gox and LDF by the throttling.

Acknowledgment Authors would like to thank to Mitsubishi Gas Chemical Co. Inc. (MGC) for supplying hydrogen peroxide used

in this studies.

FWD END AFT END

Figure 10. End surface after 7 sec. firing test.

Figure 11. Local burning surface regression sketch. Figure 12. Residue of solid fuel.


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References 1) Tsujikado, N. et al. “Experimental Studies on Air Turbo Ramjet Engines for Hypersonic Flight Vehicle (Part III)”, ISABE 2001-1237, Sept. 2-7, 2001. Bangalore, INDIA 2) Tsujikado, N. et al. “An Application of Commercial Grade Hydrogen Peroxide for Hybrid/Liquid Rocket Engine (II)” AIAA Paper #2003-5200. July 20-23 2003. 3) B. J. McBride and S. Gordon “Computer Program for Calculation of Complex Chemical Equilibrium Composition and Applications” NASA RP 1311. Lewis Research Center. June 1996. 4) M. C. Ventura and S. D. Heister “Hydrogen Peroxide as an Alternate Oxidizer for a Hybrid Rocket Booster” J. of Propulsion and Power. Vol. 11. No. 3. May-June 1995. 5) John J. Rusek “New Decomposition Catalyst and Characterization Techniques for Rocket-Grade Hydrogen Peroxide” J. of Propulsion and Power. Vol. 12. No. 3. May-June 1996. 6) Tsujikado, N. et al. “An Experimental Study of Hybrid/Liquid Rocket Engine Applied Rocket Grade Hydrogen Peroxide.” AIAA Paper #2004-3825. July 11-14 2004.

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