Hybrid Rocket Cold Flow Modeling

Wendy Cruit/EP12 205-544-1130

Due to recent renewed interest in hybrid rocket motors, MSFC has participated in several hot-fire hybrid test programs. A major concern in several of these tests has been low-frequency pressure oscillations within the motor combustion chamber. The Propulsion Laboratory's Motor Systems Division, in conjunction with Auburn University's Dr. Rhonald Jenkins, submitted a Center Director's Discretionary Fund proposal to investigate such oscillations (approved September 1995). Through ongoing research at MSFC's Structures and Dynamics Laboratory, Fluid Dynamics Division, some of the mechanisms causing the oscillations are now being studied. A laboratory-scale, two-dimensional, water-flow model of a hybrid rocket motor has been built and installed in a closed-loop, water-flow facility to aid in the research.

The hybrid model is patterned after a shear-flow water tunnel developed in 1990 to investigate fluid-flow exiting porous materials.1 The side walls are constructed of clear acrylic for flow visualization and optical velocity measurements; the top and bottom walls are constructed of a porous metal plate that simulates a solid propellant burning surface, the fuel of a hybrid motor. Oxidizer flow is also simulated with water and enters through an injector of axial or radial orientation (fig. 56). Reference figure 57 for a hybrid cold-flow model installed in a test fixture.

Flow Direction

Axial Injector

Flow Direction

111 111


Flow Direction

Radial Injector

Figure 56.—Diagram showing flow direction of axial and radial injectors.

Flow interaction between the hybrid injector and motor sections is qualitatively understood by injecting helium bubbles and recording their movement with a 1,000-frames-per-second video recorder. Currently, the flow patterns are being assessed for the different configurations in the head end and immediately downstream of the head end. The vortex shedding frequencies are also being studied to determine if there is an association between the shedding frequency and the low-frequency pressure oscillations.

Quantitative flow-field mapping is accomplished with a laser Doppler velocimeter. Velocity vectors are constructed from the two components and show the recirculation zones created by the radial injector and the stagnation regions resulting from "flame-holding" steps in the diameter between the injector and motor (fig. 58).

Figure 57.—Hybrid cold-flow model installed in test fixture.

Porous Wall

Porous Wall

Flow In

Flow Out

Porous Wall

Porous Wall \

"Flame-Holding" Step Made of Solid Material

Figure 58.—Diagram of model showing

This hybrid modeling effort is contributing to the hybrid community by generating a parametric data base of the fluid dynamics in such motors and providing test-bed capability to answer current hybrid rocket motor questions and those that will arise in the future.

'Smith, A. 1993. Porous Wall Flow Experimental Facilities. Research and Technology 1993, 178-179.

Sponsor: Center Director's Discretionary Fund flame-holding" steps.

Figure 59.—250-kilopound-force hybrid motor design concept.

Development and Demonstration of a 250-Kilopound-Force Hybrid Rocket Test-Bed

Dan M. Holt/EP12 205-544-4949

Recent increased interest in hybrid propulsion technology has prompted the formulation of an industry-led consortium to design, fabricate, and test a 250-kilopound-force thrust hybrid rocket booster system at MSFC's East Test Area. Hybrid rocket motors combine an inert solid-fuel grain with a gaseous or liquid oxidizer in an effort to realize the benefits of both liquid and solid rocket motors.

The new program, the Hybrid Propulsion Demonstration Program, combines efforts from the Hybrid Technology Option Project, the Hybrid Propulsion Technology for Launch Vehicle Boosters program, and the redirected Solid Propulsion Integrity Program. Participants include MSFC, Lockheed Martin Astronautics, Thiokol Space Operations, United Technologies Chemical Systems Division, the Rocketdyne Division of Rockwell International, Allied Signal, Lockheed Martin Manned Space Systems, and Environmental Aeroscience.

Hybrid propulsion activities have been ongoing at MSFC for several years through testing of the 11- and 24-inch motor systems at test stand 500. These motors will continue to be utilized to evaluate critical technology issues, including combustion stability, fuel utilization and web retention, nozzle and insulation materials response, and scale-up effects. Test results along with system studies will influence the large motor design. MSFC facilities, including feed, ignition, pressurization, and purge systems, will be expanded to support the testing of the 250-kilopound-force thrust motor system. MSFC science and engineering personnel will also provide for test and evaluation of the nozzle and insulation materials.

Hybrid propulsion is being considered for advanced launch vehicle applications due to advantages gained in safety, cost, environmentally benign combustion products, attractive performance, and mission flexibility relative to current rocket boosters. Recent advancements in the area of hybrid propulsion research have indicated potential application to both the expendable launch vehicle fleet and future launch vehicles. Figure 59 illustrates an early conceptual design for the 250-kilopound-force hybrid motor. The major case components include a domed forward closure, vaporization chamber, grain segment, mixing chamber, and aft closure. A submerged nozzle with exit cone is planned, with initial test capability targeted for December 1996.

The goal of the program is to demonstrate critical hybrid propulsion technologies at the 250-kilopound-force thrust scale to minimize future full-scale development risk. Preliminary concepts of the booster show promise for application to both the X-33 Advanced Technology Demonstrator and the Atlas launch vehicle. Following completion of the X-33's technology demonstration phase, hybrid boosters are being considered to provide for commercialization and mission envelope expansion. Substitution of hybrid rocket boosters on the Atlas launch vehicle increase geosynchronous transfer orbit payload capability from approximately 8.1 to more than 10 kilopound mass in projected evolutionary schemes. Figure 60 illustrates a hybrid booster concept sized to support Lockheed Martin's X-33 Advanced Technology Demonstrator and the Atlas IIAH. Detailed cost studies and booster/ launch vehicle trade studies will be performed to optimize these concepts.

Common 250-Klbf Hybrid Booster Concept

Common 250-Klbf Hybrid Booster Concept

LMC X-33H/250

Atlas IIAH

Figure 60.—Common hybrid booster concept.

Operations, United Technologies Chemical Systems Division, Rocketdyne Division of Rockwell International, Allied Signal, Lockheed Martin Manned Space Systems, and Environmental Aeroscience

Other Involvement: Advanced Research Projects Agency, U.S. Air Force Phillips Laboratory

LMC X-33H/250

Atlas IIAH

Figure 60.—Common hybrid booster concept.

Abel, T.M.; Carpenter, R.L.; Claflin, S.A.; Crawford, J.T.; and Holt, D.M. July 1995. Solid Rocket Motor Simulation and Hybrid Propulsion Testing at the Marshall Space Flight Center. Paper 95-2944, 31st American Institute of Aeronautics and Astronautics/American Society of Mechanical Engineers/Society of Automotive Engineers/American Society of Electrical Engineers Joint Propulsion Conference and Exhibit.

Sponsor: Office of Space Access and Technology

Industry Involvement: Lockheed Martin Astronautics, Thiokol Space

Technology Programs

Solid Rocket Combustion Simulation Using the Hybrid Combustion Process

Dan M. Holt/EP12 205-544-4949

Historically, test evaluation of new designs, materials, and processes for solid rocket motor nozzles and insulators at MSFC has been carried out using heavyweight, facility solid rocket motor test-beds. However, if the solid-propellant combustion environment and associated nozzle/ insulation material response is simulated using a hybrid motor test-bed, significant cost savings and expanded test flexibility could be achieved. Ongoing testing in the East Test Area, under the Large Subscale Solid Rocket Combustion Simulator program, has led to the development of a family of solid rocket combustion simulators. Testing completed to date includes subscale solid rocket combustion simulators (11 and 24

inches in diameter) that have validated the use of the hybrid combustion process for solid rocket motor simulation.

Utilizing the combustion simulator design approach—which combines an inert solid-fuel grain with a liquid-oxygen injection system in an effort to simulate a solid rocket motor, while maintaining the safety and control attributes of a liquid system—provides a unique national asset not currently available with conventional rocket motor test-beds. The combustion simulator test-beds allow for rocket motor component overtesting, margin testing, and controlled test to failure. The Large Subscale Solid Rocket Combustion Simulator program, initiated by NASA in September 1993, is being performed by Lockheed Martin Astronautics, with Thiokol Space Operations and the Rocketdyne Division of Rockwell International functioning as major subcontractors.

The 11-inch motor was used to develop regression rate correlations for the DX fuel formulation, selected to simulate the solid rocket combustion environment. Twenty-four-inch motor testing completed to date has evaluated ballistic scale-up effects, multiport motor effects, fuel overcast capability, and the initiation of a control system to enable motor operation at a constant chamber pressure and mixture ratio.

Figure 61 illustrates the various components that make up the 24-inch combustion simulator system. Testing of the 24-inch-diameter combustion simulator system was completed in July 1995 at MSFC. A total of seven tests were conducted using the circular port configuration over a wide range of motor pressures and oxidizer flow rates. Stable combustion was observed on all tests, with nozzle throat erosion ranging from 3 to 8 mils per second. This performance compares favorably with the performance of a similar-sized solid rocket motor.

In conjunction with these tests, an innovative approach for the measurement of web thickness was demonstrated. Ultrasonic transducers were installed in steel housings welded to the motor case. The development of a pulse-echo thickness measurement technique provided for real-time measurement of web thickness during motor firing. The data obtained by the ultrasonic measurement technique agree closely with pre- and posttest web thickness measurements. This system will provide a real-time estimation of fuel flow rate for use in computing mixture ratio and subsequent mixture-ratio control.

Additional testing utilized a seven-port wagon-wheel grain configuration and

Forward Closure Motor Case -S&_

Case-Bonded DX Fuel Grain _I_

n Nozzlf fm_j

Nozzle Assembly

Case-Bonded DX Fuel Grain _I_


Aft Assembly

Mixing Chamber

Figure 61.—24-inch solid rocket combustion simulator.

provided an increased understanding of the ballistic characteristics of DX fuel. Integration of a closed-loop control system—providing for realtime adjustment of an electrohydraulic valve and the corresponding adjustment in liquid-oxygen flow rate and resulting chamber pressure—was also initiated during this test series.

Figure 62 illustrates the chamber pressure profile from test L2494-090. The control system gain setting— initially set excessively high in order to respond quickly to chamber pressure deviations—resulted in an unstable control solution. However, a midtest reduction in gain provided for a stable control solution. Test L2494-090 demonstrated a significant advancement in hybrid motor control and chamber pressure (thrust) tailoring.

Testing to date on the Large Subscale Solid Rocket Combustion Simulator program has established the viability of the combustion simulator concept through testing with the 11- and 24-inch solid rocket simulators. A valuable data base has been compiled to allow for future evaluation of advanced solid rocket motor materials and processes. The overall system will provide invaluable advancement of solid propulsion.

Abel, T.M.; Boardman, T.A.; and Crawford, J.T. November 15-17, 1994. Rocket Nozzle Materials Testing Using a Solid Rocket Combustion Simulator. Joint Army, Navy, NASA, and Air Force Rocket Nozzle Technology Subcommittee Meeting, Seattle, Washington.

Sponsor: Office of Space Access and Technology

Industry Involvement: Lockheed Martin Astronautics, Thiokol Space Operations, Rocketdyne Division of Rockwell International o


Figure 62.—Control system integration test.

Figure 62.—Control system integration test.

Solar Thermal Propulsion Thrusters and Cryogenic Fluid Management

Leon J. Hastings/EP25 205-544-5434

A number of technologies must mature before full-scale development of the solar thermal propulsion concept can be undertaken. These include the absorber/thruster and the subcritical liquid-hydrogen storage/ feed system. Basically, the solar propulsion concept involves focusing sunlight on a high-temperature absorber that heats hydrogen to 2,500 Kelvin and expels it from a nozzle for thrust in the 4.4- to 44.0-Newton range at high specific impulse (800 to 900 seconds). The MSFC effort consists of three basic elements:

(1) development of absorber/thruster fabrication and assembly techniques,

(2) development of an MSFC solar test facility, and (3) testing of a cryogenic fluid management subsystem. Activities and accomplishments for each of the three elements are summarized below. Reference figures 63 and 64.

Absorber/thruster: High-temperature absorber/thruster operation presents significant technology challenges involving heat transfer and materials selection and design. As presently envisioned, the absorber/thruster will be constructed primarily of a tungsten/ rhenium-alloy inner and outer shell surrounded by a carbon foam or graphite insulation. The gaseous hydrogen will flow through the passage between two shells, absorb the energy focused into the inside cavity,

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