(c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

A01-34045

AIAA 2001-3244

Development and Lab-Scale Testing of aGas Generator Hybrid Fuel in Support ofthe Hydrogen Peroxide Hybrid Upper StageProgram

Gary K. Lund, Wm. David Starrett and Kent C. JensenATK, Thiokol Propulsion Corp.Brigham City, Utah

37th AIAA/ASME/SAE/ASEEJoint Propulsion Conference and Exhibit

8-11 July 2001Salt Lake City, Utah

For permission to copy or to republish, contact the copyright owner named on the first page.For AIAA-held copyright, write to AIAA Permissions Department,

1801 Alexander Bell Drive, Suite 500, Reston, VA, 20191-4344.

(c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

AIAA 2001-3244

DEVELOPMENT AND LAB-SCALE TESTING OF A GAS GENERATORHYBRID FUEL IN SUPPORT OF THE HYDROGEN PEROXIDE HYBRID

UPPER STAGE PROGRAMGary K. Lund, Wm. David Starrett and Kent C. Jensen

ATK, Thiokol Propulsion Corp.Brigham City, Utah

Abstract

As part of a NASA funded contract to develop anddemonstrate a gas generator cycle hybrid rocket motorfor upper stage space motor applications, thedevelopment and demonstration of a low sensitivity,high performance fuel composition was undertaken.The ultimate goal of the development program was todemonstrate successful hybrid operation (start, stop,throttling) of the fuel with high concentration (90+%)hydrogen peroxide. The formulation development andlab-scale testing of a simple DOT Class 1.4c gasgenerator propellant is described.

Both forward injected center perforated and aftinjected end burner hybrid combustion behavior wereevaluated with gaseous oxygen and catalyticallydecomposed 90% hydrogen peroxide. Cross flow andstatic environments were found to yield profoundlydifferent combustion behaviors, which were furthergoverned by binder type, oxidizer level and,significantly, oxidizer particle size.

Primary extinguishment was accomplished viamanipulation of PDL behavior and oxidizer turndown,which is enhanced with the hydrogen peroxide system.Laboratory scale combustor results compared verywell with 11-inch and 24-inch sub-scale test resultswith 90% hydrogen peroxide.

Introduction

Growing international competition for delivery oforbital payloads has prompted increased governmentand industry investment in advanced propulsionsystems. NASA is funding efforts to investigateperoxide based propulsion systems, including hybrids.

Hybrid rockets have long been considered a low costalternative due to their inherent safety, storability, andoperational efficiency. Combining these features withthe clean burning, storable, non-toxic attributes ofhydrogen peroxide yields an excellent combination.

NASA has recently funded an effort to develop aperoxide hybrid upper stage propulsion system. Theteam, which is comprised of Lockheed Martin,Thiokol, and Rocketdyne, seeks to design, integrate,and test a peroxide thrust chamber assembly that willbe capable of delivering 10,000 pounds of thrust at avacuum specific impulse of 310 seconds forapproximately 360 seconds.

The candidate system chosen for investigation is a gasgenerator design incorporating a partially oxidized,fuel-rich propellant with combustion initiated by, andmade more efficient with, the injection of highconcentration hydrogen peroxide. Multiple starts,steady state operation, and extinguishment of thehybrid motors are planned as part of this investigation.This paper will focus on the development, lab-scaletesting/evaluation, and selection of the fuelformulation for this system.

Technical Approach

The fuel requirements for the gas generator cyclehybrid are challenging from a formulation perspective.Specifically, the fuel is to function as a propellant withrespect to combustion characteristics (i.e. exhibitpressure dominated burn rate), yet readily extinguishupon oxidizer flow termination. Fuel formulationgoals for this program were as follows:

Copyright © 2001 ATK, Thiokol Propulsion Corp.Published by the American Institute ofAeronautics and Astronautics, Inc., with permission.

1American Institute of Aeronautics and Astronautics

(c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

1) Vacuum specific impulse with 90%hydrogen peroxide (H2O2), 200 psiachamber pressure, 100/1 expansion ratio;> 310 seconds.

2) Available, low cost ingredients.3) Low hazards; DOT Class 1.4c.4) Moderate regression rate: ~ 0.1 inches per

second @ 200 psia.

In addition, the fuel had to exhibit efficient, stablecombustion at operating pressures in the absence ofexternal oxidizer. To this end, some emphasis wasplaced on minimizing soot and char production duringthe propellant fuel combustion. The approach chosenutilized the improved combustion characteristics ofpolyether (PPG) based binders with the conventionaloxidizer, ammonium perchlorate (AP). Figure 1illustrates the theoretical Isp performance of severaloptions with 90% peroxide. This illustrates that fairlyconstant Isp performance is achieved at optimumstoichiometry for a fairly wide range of fuel types andAP loadings, albeit at substantially differentoxidizer/fuel ratio's (O/F's). As AP content and/orPPG content increases, optimum O/F shifts to lowervalues.

Isp vs Mix Ratio With 90% Peroxide

§.0> 300

i5 280Q.

8CM

« 260

O/F ratio

Figure 1. Theoretical Isp vs. mix ratio for severalfuel variations with 90% hydrogen peroxide.

Fuel Formulation and Testing

Fuel formulation screening was conducted with curedstrands, small ballistic test motors, and with a small-scale laboratory hybrid combustor operated withgaseous oxygen (GOX) or 90% hydrogen peroxide.AP was selected as the oxidizer. AP levels andparticle size were varied Binders examined includedhydroxy terminated polybutadiene (HTPB), acommercial PPG, or mixtures of the two.

In general, it was found that compositions utilizingHTPB as the sole binder tended to produce largeamounts of soot during combustion. This behavior ledto concern over ultimate combustion efficiency withthis approach. In addition, low pressure deflagrationwith HTPB based propellants tended to be verypersistent, which is anticipated to interfere withextinguishment of the system. PPG tended to exhibit ahigher pressure deflagration limit (PDL) than HTPB.

On the other hand, HTPB possesses a more favorablefuel value than PPG, providing improved Isp andhigher optimum O/F. HTPB based formulations alsoprovided much higher burn (regression) rates thanPPG alone. A blended binder approach was, therefore,pursued as a compromise.

Figure 2 illustrates cured strand ballistic results for atypical formulation. The presence of the plateau isinteresting in that it seemed to be characteristic ofthese fuel rich compositions and persisted in the smallmotor ballistic test results as well.

2

s1ff

1

Closed Bomb Strand Burning RatePPG Binder

6 7 8 91Q2 2 3 4 5 6 7 8 S-jgS 2

Pressure (psi)

Figure 2. Closed bomb strand burning rates forPPG/AP gas generator propellant

In general, obtaining ballistic data with thesecompositions proved challenging due to difficultignition, extinguishment and L* sensitivity. Typicaltest-to-test variability was found to be on the order of15% to 20% in the absence of external oxidizer. Duringthese experiments, it was found that AP particle size hada significant effect on ignition and sustained burningunder these conditions. The presence of carbon black inthe formulation also affected the ignition characteristics.Many of the fuel formulation candidates were also testedin a small-scale laboratory hybrid combustor based onthe ballistic test motor. Grains were either a 1.5-inch or2-inch diameter center perforated (CP) configurationwith almost identical port diameters. As illustrated in

American Institute of Aeronautics and Astronautics

(c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

Figure 3, oxidizer (typically GOX) was injected into theforward end of the motor during operatioa

Figure 3. Schematic of 2-inch diameter lab-scalehybrid combustor.

Variables evaluated in this device consisted of oxidizermass flow (flux), pressure and formulation. In thesetests, ignition was accomplished by conventional squib,and the test terminated by oxidizer cut-off. Regressionrates were determined by measuring the change in webthickness post-test or, alternatively, via weight loss. Themajority of the testing with the CP configuration wasdone with the 1.5-inch diameter motor, and the resultsindicated that, for most formulations tested, regressionrate was very much governed by pressure with little ifany oxidizer flux response noted as illustrated in Figure4.

Tests 3282 - 3290

Figure 4. Regression rate vs. oxygen mass flux andchamber pressure for baseline fuel.

The (CP) grain configuration was operated with 90%hydrogen peroxide in addition to oxygen. In this case, aportable test stand and catalyst system manufactured byGeneral Kinetics, LLC was interfaced with thecombustor. Figure 5 is a photo of this system.

Figure 5. The hydrogen peroxide feed system andlab-scale combustor.

The 2-inch diameter CP grain configuration was utilizedfor the peroxide tests, principally to provide additionalweb for extended test times. Squib ignition was notnecessary with the CP configuration since the hot,decomposed peroxide proved adequate for hypergolicignition. Figure 6 is a reproduction of a typical CP testwith this configuration showing the peroxidemonopropellant operation phase and the grain ignition.

Test 3492 Motor Only Purge, 0.247 Nozzle

Time (sec)

Figure 6.2-inch CP test with hydrogen peroxide.

Operation was reasonably stable and the grainextinguished readily. The O/F ratio for this test was 2.6which is a little higher than stoichiometric (at roughly2.0). In general, the regression rates measured with theperoxide were comparable to those obtained with GOX,particularly at lower pressures. Figure 7 comparesregression rates as a function of pressure for the TU-628(1.5-inch) and TU-172 (2-inch) motors with GOX andhydrogen peroxide respectively.

American Institute of Aeronautics and Astronautics

(c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

B£ n nf-

&0.04 -

(

CR» w/ Various Oxidizer

*• '.

• :•Go«TU>28CP

+ « !* • •

+ ,, / i H20Z/nm« 4 i CP-Good

) 100 200 300 400 500 600 700Max Chamber Pressure (psi)

Figure 7. Gas Generator CP Performance with GOXand H2O2.

Further experimentation revealed that the high pressuredeviation observed in the plot was most likely due to therelative sizes of the motors and the presence of somevelocity related flow effects inherent in the smaller 1.5-inch CP motor.

Development of an end-burning configuration for thelab-scale motor allowed further exploration with regardto flow environment. This motor, illustrated in Figure 8,utilized a 2-inch diameter grain and could be operated ata variety of length to diameter (L/D) and L* (ratio ofchamber volume divided by throat area) conditions bysimply changing the length of the primary cylindricalcombustion chamber. The end-burner combustor allowedflow effects during combustion to be varied by alteringthe angle of impingement of the injector port and thedistance from injector to fuel grain surface.Environments ranging from cross-flow to directlyimpinging to minimal flow (static) were tested.

Figure 8. End burner configuration at moderate L*(1000 - 2000 in.) chamber length illustrating angledand radial flow geometry options.

As with the CP grain tests, both GOX and 90% hydrogenperoxide were used for fuel testing in the end burner.The results from these tests led to a few surprises. As

shown by the pressure plots in Figures 9 and 10,combustion chamber L/D, and therefore L*, had asignificant effect on combustion performance with somefuels. Both tests shown were conducted with identicalfuel, GOX mass flow rate and throat area, differing onlyin length of the combustion chamber (3.5 inches versus7.5 inches).

Test 3420 3.5" Tube, 0.144 Nozzle

Time (sec)

Test 3421 7.5" Tube, 0.144 Nozzle

X-N 50°

1g> 400

3tf>£ s o o

Time (sec)

Figures 9 and 10. Pressure vs. time for end burner at3.5 inch and 7.5 inch combustion chamber lengths.Pressure spike at 0.5 seconds is from the squib.

Based on these results, it became apparent that someformulations that performed well in a CP cross-flowenvironment were unduly sensitive in the more static-flow end burner environment Not only were thechamber pressures reduced at large L/D, but regressionrates were halved as well. This observation implied thatthe combustion was not stable under conditions of verylow oxidizer flow, completely unacceptable for propermotor operation. Fortunately, the formulation respondedwell to an AP size adjustment, which proved to be an

American Institute of Aeronautics and Astronautics

(c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

effective remedy. As shown in Figure 11 (compared toFigure 9), identical motor pressure and regression rateswere achieved at both L/D conditions with the modifiedfuel. The final formulation was designated DL-H485and was further characterized in the end burner with90% hydrogen peroxide.

Test 3543, 7.5" Tube, 0.135 Nozzle

Time (sec)

Figure 11. Long L/D GOX test with improvedformulation.

The lack of significant oxidizer cross-flow orimpingement on the fuel grain surface led todramatically increased ignition delays with H2O2relative to those observed with the CP tests.Consequently, squib ignition was required for these tests,coupled with adequate delay time for sufficient heatingof the hardware (typically 15 seconds).

Regression rates obtained with the baseline fuel withboth GOX and peroxide in the end burner are comparedin Figure 12. In this case, the regression rates are againvery similar with those from the peroxide tests beingslightly higher than with the GOX. This is suggested tobe due to the fact that the O/F ratios for all of the testswas roughly 1.5 to 3, which is fairly close to optimumfor 90% peroxide but considerably oxidizer rich for theGOX tests. Since this equates to lower flametemperature, it is not unreasonable to expect some effecton observed regression rate. Overall, a regression rate ofabout 0.06 inches per second at 200 psi with a pressureexponent of 0.3 was obtained for the DL-H485 fuel and90% hydrogen peroxide.

End Burners w/ Various Oxidizer

0.08 -i

0.07-

f\ c\a

? 0.05 -

1 0.04 -

k 0.03 -

1 0.02 -CQ

0.01 -

• :

_^ -̂"*"^^?^^

j/^s^Sy*

\ H202<250psi BGox |

o oo - —————— , ————— r ——— , —————— ,0 100 200 300 400

Avg Max Chamber Pressure (psi)

Figure 12. Regression rate comparison for DL-H485fuel end burner with GOX and 90% peroxide.

Variation of the oxidizer injection angle and distance tofuel surface were performed to confirm the sensitivity ofthe fuel to flow effects at various extremes. As shown infigure 13, an increase in regression rate was observedwhen impinging flow prevailed in the combustor. Theseregression rates are similar to what was observed in theCP tests and also tends to imply that the fuel is sensitiveto flow (flux) effects. This may appear contrary to thedata presented in Figure 4 which suggests minorsensitivity over the ranges tested. These rangesrepresent modest velocity conditions (all at fairly highoxygen mass flux) as opposed to the conditions in theend burner. Under these conditions, the effects between adirect impinging flow and a non-impinging condition areexpected to be considerably greater.

Injector Flow EffectsLab Scale 2-inch End Burner - DL-H485 IXiel - Gox

0.160.14 -0.12 -0.10 -0.08 -0.060.040.02 -0.00

100 200 300 400 500 600Max Chamber Pressure (psi)|* Angbd2" "Radial2" |

700

Figure 13. Lab-scale Injector Flow Effects.

American Institute of Aeronautics and Astronautics

(c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

Scale-Up

DL-H485 was successfully scaled up to the 5-gallon mixlevel (50 Ib.) to support 11-inch and 24-inch diameterstatic motor testing. Grains were fabricated as endburners by casting into a lined silica-phenolic "cup" andcured. TTiese were then cartridge loaded into therespective test hardware for testing. Figure 14 is aphotograph of a representative 11-inch fuel grain.

Summary and Conclusions

A gas generating fuel formulation (DL-H485) has beendeveloped which meets the goals of the program, i.e.,theoretical Isp>310 seconds, low cost ingredients, lowhazards, moderate regression rate, self-deflagrating,stable and efficient combustion in both low and high L*configurations, and which extinguishes on oxidizershutoff.

The lab-scale combustor has proved to be invaluable inthe development of gas generator hybrid fuelformulations. The results from lab-scale testing closelyagree with the larger 11-inch and 24-inch motorsindicating excellent scale-up characteristics. This isobviously quite advantageous for future hybrid fuel andmotor development efforts.

Fuel formulation and testing to date have demonstratedthe viability of a peroxide hybrid upper stage motorsystem.

Figure 14. As-cast 11-inch solid fuel grain.

Testing of the 11-inch and 24-inch motors wasconducted at Stennis Space Flight Center with 90%hydrogen peroxide. Test operating conditions weresimilar to those employed in the lab-scale studies withrespect to test times, pressures, O/F and L*. The 11- and24-inch test results agreed closely with the lab-scaleresults discussed herein.

Hazard testing of DL-H485 resulted in it's being granteda DOT Class 1.4c hazards classification which hassignificant ramifications for manufacture, shipping,storage and handling. This allows the gas generator cyclehybrid motor as developed here to retain one of the moresignificant advantages of classical hybrid designs - thatof reduced handling and storage hazards relative toconventional propellants.

American Institute of Aeronautics and Astronautics