Development of Thermal Barriers for Solid Rocket Motor Nozzle Joints
Bruce M. Steinetz
Glenn Research Center, Cleveland, OhioPatrick H. Dunlap, Jr.
Modern Technologies, Corp., Middleburg Heights, OhioJune 1999
The Space Shuttle solid rocket motor case assembly joints are sealed using conventional O-ring seals. The 5500+°F combustion gases are kept a safe distance away from the seals by thick layers of insulation. Special joint-fill compounds are used to fill the joints in the insulation to prevent a direct flowpath to the seals. On a number of occasions, NASA has observed in several of the rocket nozzle assembly joints hot gas penetration through defects in the joint-fill compound. The current nozzle-to-case joint design incorporates primary, secondary and wiper (inner-most) O-rings and polysulfide joint-fill compound. In the current design, 1 out of 7 motors experience hot gas to the wiper O-ring. Though the condition does not threaten motor safety, evidence of hot gas to the wiper O-ring results in extensive reviews before resuming flight. NASA and solid rocket motor manufacturer Thiokol are working to improve the nozzle-to-case joint design by implementing a more reliable J-leg design and a thermal barrier.
This paper presents burn-resistance, temperature drop, flow, and resiliency test results for several types of NASA braided carbon-fiber thermal barriers. Burn tests were per-formed to determine the time to burn through each of the thermal barriers when exposed to the flame of an oxy-acetylene torch (5500 °F), representative of the 5500 °F solid rocket motor combustion temperatures. Thermal barriers braided out of carbon fibers endured the flame for over 6 min, three times longer than the solid rocket motor burn time. Tests were performed on two thermal barrier braid architectures, denoted Carbon-3 and Carbon-6, to measure the temperature drop across and along the barrier in a compressed state when subjected to the flame of an oxyacetylene torch. Carbon-3 and Carbon-6 thermal barriers were excellent insulators causing temperature drops through their diameter from 2500 to 2800 °F. Gas temperatures 1/4" downstream of the thermal barrier were within the downstream Viton O-ring temperature limit of 600 °F. Carbon-6 performed extremely well in subscale rocket "char" motor tests when subjected to hot gas at 3200 °F for an 11-sec. rocket firing, simulating the maximum downstream joint cavity fill time. The thermal barrier reduced the incoming hot gas temperature by 2200 °F in an intentionally oversized gap defect, spread the incoming jet flow, and blocked hot slag, thereby offering protection to the downstream O-rings.