TRW 40 klbf LOX/RP-1 low cost pintle engine test results

Abstract

One of NASA's goals is to develop and demonstrate next-generation technology that will enable industry to provide truly affordable and reliable access to space. Their mission is to provide the necessary technology to reduce dramatically both the cost of placing payloads in space and the cost of in-space transportation. This includes the next generation manned and unmanned shuttle and the heavy lift launch vehicles for future Mars and space exploration. History shows that about 40% of the vehicle cost is the cost of the rocket engines. A low-cost rocket engine could dramatically reduce the overall vehicle cost For the last 30 years, TRW has focused on low cost, simplicity and reliability in the design of the TRW single element coaxial pintle injector rocket engine. Early work at TRW involving the LOX/Kerosene propellants started with a 2K engine, tested in the late 1960s, achieving 96% combustion efficiency. In the early 1970s, the TRW Holloman sled engine was retrofitted to burn LOX/RP-1 at 50 kfbf thrust. This engine achieved 93% combustion efficiency. Also, stable combustion was demonstrated in bomb tests. In the early 1990s, TRW tested an LOX/LH2 engine at both the 16.5 and 40K thrust levels, using the same basic injector hardware. In 1995, this engine was retrofitted with injection rings sized for LOX/RP-1 and tested at 13K thrust. This engine demonstrated 98% combustion efficiency in runs with an ablative chamber and was also shown to be stable after bomb testing. Recently, TRW tested a pressure fed 40K LOX/RP-1 low cost pintle engine (LCPE) and the test results are reported herein. *Project Engineer, Member of AlAA **Chief Engineer, Member of AiAA INTRODUCTION SUMMARY In a test program co-sponsored by TRW and the Air Force, TRW tested a pressure-fed 40K LOX/RP-1 engine at the Energetic Materials Research and Testing Center (EMRTC) Rocket Test Site in Socorro New Mexico in November and December of 1999. Testing at 25 klbf thrust and 40 klbf thrust were accomplished, using one injector and chamber assembly with only changes to the injection orifice elements, fn all, 15 hot fire tests were conducted, six at 25 klbf thrust level, and nine at 40 kibf thrust level, demonstrating combustion efficiencies of up to 98.4% of theoretical C* (ODE). Fuel film cooling of the chamber was performed at the 40K thrust level in 4 separate tests, with fuel film cooling flow rates of 4%, 6% and 9% of the total fuel flow demonstrated. Test durations were one to five seconds using a copper heat-sink chamber on loan from NASA Glen Research Center. The 25 klbf tests were operated at 250 psia chamber pressure and a target mixture ratio of 2.4, while the 40 klbf tests were operated at 390 psia chamber pressure and a mixture ratio of 2.25. Data was gathered on seven different oxidizer slot geometries, and three different fuel injection velocities, providing basic information needed for optimization of engine performance. HARDWARE DESCRIPTION The injector and chamber for this test series (Figure 1) were originally built in the early 1990's for testing of the 16 kibf and 40 klbf LOX/LH2, and 13K LOX/RP-1 at NASA Glen (then NASA LeRC), described in References 1 through 4. Modifications for this program involved replacing the original 4-inch pintle diameter with a new 5.25 Copyright © 2000 by TRW Inc. Published by the American institute of Aeronautics and Astronautics, Inc., with permission. American Institute of Aeronautics and Astronautics (c)2000 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.