Development of Thermal Barriers for Solid Rocket Motor Nozzle Joints

AIAA-99-2823 NASA/TM--1999-209278 Development Solid Rocket Bruce M. Steinetz Glenn Research Patrick Center, H. Dunlap, Modern of Thermal Barriers Motor Nozzle Joints Cleveland, Jr. Technologies Prepared for the 35th Joint Propulsion Corporation, Conference cosponsored by the AIAA, Los Angeles, California, National Space Aeronautics Administration Glenn Research June 1999 Ohio Center and ASME, June and Middleburg 20-24, Heights, Exhibit SAE, 1999 and ASEE Ohio for Acknowledgments The authors Bruce Bond's greatly acknowledge Doug Frost's and Steve Hicken's (Thiokol) thermal (Albany-Techniweave) assistance in fabricating the thermal barriers; and Lauren Yee's (NASA) assistance in test support. barrier test results; and Tom Doeberling's Trade names or manufacturers' names are used in this report for identification only. This usage does not constitute an official endorsement, either Aeronautics expressed or implied, by the National and Space Administration. Available NASA Center 7121 Standard for Aerospace Drive Hanover, MD 21076 Price Code: A03 Information from National Technical Information Service 5285 Port Royal Road Springfield, VA 22100 Price Code: A03 AIAA-99-2823 DEVELOPMENT OF THERMAL BARRIERS FOR SOLID MOTOR NOZZLE JOINTS Bruce National M. Steinetz* Aeronaulics and Space Glenn Research Cleveland, Administration Center Ohio 44135 H. Dunlap, Patrick Modern Technologies Middleburg Jr.; Corporation Heights, Abstract Ohio 600 44130 °F. Carbon-6 rocket The joints Shuttle Space sealed are 5500+°F solid using combustion from the seals motor casc ()-ring conventional gases by thick fill compounds rocket seals. layers of insulation. The arc used to fill the joints Special assembly in the joint-fill design joints hot gas penetration The current incorporates most) O-rings current motor I out to ()-ring. safety, Though evidence in extensive reviews solid motor rocket improve reliable heft)re resuming manufacturer braided carbon-fiber tormed to determine thermal barriers acetylene thermal (5500 °F), types and to representative seals research drop, of NASA of the of an oxy- of the 5500 burn time. braid architectures, measure Tests were performed denoted the temperature in a compressed state oxyacetylene torch. barriers excellent drops were through temperatures within their drop across when and insulators diameter Viton from °F motor thermal barrier and Carbon-6, Carbon-6 causing seals do not jetted tothe arc found °F. barrier Gas several thereby materials transient seal designs that seal as compliant and pressure require- 3 showed that these a few seconds Shuttle Solid a thermal de- Previous when rocket to evolve barrier motor. the braided for sub- gases use that Thus, rope at extreme temperatures. The Space Shuttle joints 5500+ reusable are scaled the seals solid rocket motor usingconventional °F combustion from insulation. than led to of newly wires. and Dunlap were considered into out rope °F) com- designs hot 55(X)+ °Fcornbustion in the Space design braided temperature last for more seal engine braided Steinetz gases by thick Special.joint-fill are kept layers of phenolic compounds seals. two stages of protection a direct llowpath reaching the seals. charring due to parasitic leakage seals paths to ['ill the l]ow-patb of the 900-psi Occasionally, or rubber are used a direct these seals. a safe distance to prevent Normally, (RS RM) O-ring .joints in the insulation to the are enough hot gases have from experienced thai open up in the were limit of Branch, Copyright © 1999 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17, U.S. Code. The U.S. Government has a royahy-free license to exercise all rights under the copyright claimed herein for Governmental Purposes, All other rights arc reserved by the copyright owner. 1 American by spread slag, to both seal and serve extremely other to prevent thermal temperature *Senior Research Engineer. Mechanical Components Member AIAA. Resean:h Emgineer. of an temperature 25(_) to 2800 hot ( 1500 to 2000 seals under aggressive away to the barrier to the flame of the thermal ()-ring rocket and along subjected Carbon-3 1/4" downstream the downstream on two Carbon-3 thermal temperature O-rings. and supcralloy the ability 1,2 However, The than the solid blocked gas turbine fibers yielded over 6 rain, longer and The at the gap defect, to the downstream of rope demonstrated assembly times oversized in advanced ceramic solid rocket motor combustion temperatures. Thermal barriers braided out of carbon fibers endured the flame tbr three gas hot gas simulating fill time. cavity The need tor high temperature pliant ments. each flame firing, in subscale to hot Introduction mounts tests were per- through to the rocket incoming flow, protection veloped barrier. Burn to burn exposed NASA temperature barriers. the time results are working for several ,jet incoming offering subjected joint the the development by implementing burn-resistance, when torch flight. and a thermal test results to not threaten ()-ring Thiokol joint design design presents and resiliency does the In the hot gas of hot gas to the wiper J-leg This paper flow, the condition the nozzle-to-case a more joint (inner- experience downstream reduced well when defects compound. of 7 motors l 1-see. 22(X) °F in an intentionally and wiper secondary maxinmm an joint- nozzle-to-case and polysulfidejoint-fill design, the wiper primary, through fl)r extremely tests barrier in the insulation compound. °F perlormed motor away prevent a direct flowpath to the seals. On a number of occasions, NASA has observed in several of the rocket nozzle "char" 3200 assembly are kept a safe distance ROCKET Institute of Aeronautics and Astronautics joint-fill compounds during rocket operation. Inspection during disassembly of Space Shuttle solid rocket motor nozzle joints from RSRM-44 and RSRM-45 revealed ()-ring erosion of Joint 3 primary O-ring seals 4 (Fig. 1). Subsequent improvements in joint-fill compound application-techniques have apparently overcome the Joint 3 charring problem. However, a number of nozzle joints including the nozzle-to-case joint and Joint 2 continue to show hot gas penetration through the joint fill compound. The current nozzle-to-case joint design incorporates primary, secondary and wiper (innermost) ()-rings and polysulfidc joint-fill compound. In the current design, I out of 7 motors experience hot gas to the wiper ()-ring. Though the condition does not threaten motor safety, evidence of hot gas to the wiper O-ring results in extensive reviews belore 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, and eliminate the jointfill compound (Fig. I). The J-leg is molded into the insulation and contacts the mating surface of the adjoining element. Rocket pressurization acts to further preload the J-leg increasing its effectiveness. The basic J-leg design has been applied successfully to fixing the field joints in the redesign el'fl+rt lollowing the Challenger accident. s The thermal barrier, compressed between the J-leg and adjoining clement, is intended to resist any hot gases the J-leg does not block and prevent them from reaching the wiper ()-ring. The braided carbon thermal barrier being developed at NASA Glenn is the leading candidate based on the results presented herein. The thermal barrier for the Shuttle solid rocket motor has unique requirements, others: including the following, amongst I. Sustain extreme temperatures (2500 to 55(X) °F) during solid rocket motor burn (2 min and 4 see.) without loss of integrity. 2. Drop incoming gas temperatures (up to 3200 °F) in the joint to levels acceptable to Viton O-rings (<600 °F, short-term) to prevent ()-ring damage+ including char and erosion. 3. Exhibit some permeability to permit the joint cavity (between thermal barrier and O-ring) to reach chamber pressure (900-psi) in acceptable time. 4. Exhibit adequate resiliency/springback to accommcx:late limited joint movement and manufacturing tolerances in these large (8.5 It. diam.) nozzle segments. 5. Diffuse/spread incoming narrow (0.08 in. diam.) hot gas jets to reduce their damaging effects on the downstream ()-rings. 6. Block hot slag (i.e., molten alumina, etc.) entrained in gas stream from reaching O-rings. American Institute Steinetz and Dunlap 3 pertormed a number of tests on 0.125- and 0.2(X)-in. diameter braided carbon-fiber thermal barriers demonstrating that they met the burn-resistance, permeability, and resiliency criteria. The main objective of the current study is to fully characterize two braided carbon fiber thermal barrier designs (denoted Carbon-3 and Carbon-6) by assessing their transient thermal response when subjected to a high temperature torch and by characterizing their permeability, resiliency, and burn-resistance. Thc Carbon-6 design is currently being tested by both NASA and Thiokol for the nozzle-to-case joints of the Shuttle solid rocket motor. Subscale rocket "char" motor tests were performed to assess the thermal barrier's (Carbon-6) thermal response and heat resistance under actual rocket conditions. Test Apparatus Thermal Barrier and Procedures Specimens Carbon-3 and Carbon-6 were subjected to burn, temperature drop, flow, and compression tests. Carbon-6 was also tested in a subscale char motor. Limited testing was performed on the Carbon-4 design. Table I summarizes the relevant architecture parameters for the thermal barrier designs that were tested. All thermal barriers wcrc composed ofa uniaxial corc of fibers overbraided with various numbers of sheath layers. The Carbon-6 design had ten sheath layers and a 0.26-in. diameter. Carbon-6 had good flexibility and compliance properties because it was braided with a more open architecture. The Carbon-3 design had a 0.20-in. diameter and was made with a large degree of uniaxial core fibers overbraidcd with five sheath layers. Carbon-3 was a tight braid that was not as flexible as Carbon-6. Carbon-4 had 4.4×10 -4 in. (I I Jam) pitch-based Amoco P25 fibers in its core to evaluate core fiber diameter effects on performance, while the core fibers of all the other carbon thcrmal barriers were 2.76×10 -4 in. (6.9 Jam) PAN-based Grafil type 34700 fibers. PAN-based Thornel T-300 carbon fibers with a 2.8× 10-4 in. (7 _m) diameter all the thermal Thermal Barrier were used in the sheaths of barrier designs. Porosity Measurements To assess thermal barrier porosity while under compression, samples of the Carbon-3 and Carbon-6 designs were examined in a compressed state using a photographic stereomicroscope. Four !/2-in. long specimens of I'x)th types of thermal barriers were prepared and weighed using a precision electronic balance. The exact length of each specimen was measured using vernier calipers. Each specimen was then clamped between two steel plates and subjected to a 2()_: compression. While the specimens were compressed, a light layer of cyanoacrylic glue was applied to the surface 2 of Aeronautics and Astronautics of eachspecimen so thattheywouldmaintaintheir compressed shape uponremoval fromthefixture. Fourspecimens wereexamined forbothCarbon-3 andCarbon-6. Bothends ofeach specimen were examined andphotographed atI0Xinthemicroscope sothateight cross section photos were examined for both thermal barrier designs. Each cross section assumed an ellipsoidal shape in its compressed state. The dimensions of each ellipse were measured using vernier calipers. These dimensions were then used to calculate the cross sectional . area of both ends of each specimen. An average cross sectional area was calculated tor each specimen and multiplied by the specimen length to determine the specimen volume. Specimen density was then calculated by dividing the weight of the specimen by its volume. An average density at 20% compression was found for both Carbon-3 and Carbon-6 by averaging the densities of the four specimens of each design. The porosity of each thermal barrier design at 20% compression was calculated using the following Porosity relationship: = I- (gThe,',,,al Barrier/PCarbon Fiber) In this relationship, the density of each thermal barrier design was divided by the density of an individual carbon fiber (0.064 Ib/cu.in.). Thus, a thermal barrier design would have a porosity of zero if it had no gaps and assumed the density of an individual fiber. Burn Tests A screening test was developed to evaluate thermal barrier burn resistance under sire ulated rocket motor combustion temperatures (5500 °F) by aiming a "neutral" flame 6 of an oxyacetylene welding torch at the center section of a 4-in. thermal barrier specimen. In these tests, the amount of time required to completely cut through the specimen measured was measured. Time for cut-through was from the instant the flame touched the specimen until the specimen was completely cut into two separate pieces. A detailed description and an illustration of the fixture used to perform these tests can be found in the paper by Steinetz and Dunlap. 3 Temperature Drop Tests A test fixture was developed to measure the temper- ature drop across and along the thermal barriers in a compressed state when subjected to the neutral flame of an oxyacetylene torch simulating rocket temperatures (Fig. 2). Flow was drawn through the thermal barrier using a vacuum roughing pump to lower pressure on the downstream side of the thermal barrier while leaving the upstream side at ambient conditions. Flow through the thermal barrier was measured using a flow meter positioned between the fixture and the roughing pump. The volume American Institute downstream of the thermal barrier was an enclosed plenum chamber sealed by an O-ri ng between the bottom plate and a top plate. The thermal barrier was compressed at 20% linear compression. Other compressions are possible by placing shims under the thermal barrier. The fixture was made out of phenolic insulation having low thermal conductivity that simulates the solid rocket motor insulating material and minimizes parasitic heat loss. The torch flame was applied to the thermal barrier to simulate a leak path of hot gases through the nozzle joint. The flame with temperatures up to 32(XJ°F was positioned on a small area of the thermal barrier. An "'iris plate" with a 0.084-in. diameter hole concentrated a "laser-like" column of flame onto the thermal barrier, simulating a hot gas jet flowing through the rocket nozzle joint. The iris plate was positioned about I/4-in. away from the specimen. The jet was directed at the center of the specimen both span- and height-wise. To measure the surface temperature distribution along the thermal barrier span, thermocouples were placed on both the upstream (hot)and downstream (cold) sides. The thermal barrier specimen sat between these two rows of thcrmocouples in a 0.040-in.-deep groove. The thermocouples measured how the flame spread along the thermal barrier, how much temperature drop occurred across the thernnal barrier, and how heat was conducted along its length. The fixture was instrumented with seven thermocouples upstream of the thermal barrier and eight downstream Ihermocouples. On the upstream side, the center Type B thermocouple was placed directly in line with the center of the hole in the iris plate so that it measured the hottest flame temperature at the surface of the thermal barrier. Type B thermocouplcs were then positioned I/4-in. on either side of the center thcrmocouple (Fig. 2). The remaining four thermocouplcs on the hot side were Type K thermocouples, and they wcrc placed 1/2 and I in. on either side of the center thermocouple. Seven of the eight Type K thermocouples downstream of the thermal barrier were spaced so that they were directly in line with those upstream of the thermal barrier. The remaining Type K thermocouple was positioned I/4 in. (approxinmtely one thermal barrier diameter) downstream of the thermal barrier in line with the center thermocouple and measured the bulk air temperature. Thermocouple selection. Fine gage wire open-bead thermocouples were used to quickly and accurately measure changes in the surface temperature distribution along the thermal barrier. The time constant and response rate of a thermocouple is controlled by the size of its wires and the diameter of the junction ball that is lbrmed between the wires. The wire diameters used lot the Type B and Type K thermocouples were 0.010 and 0.0125 in., respectively. A typical thcrmocouple junction ball has a diameter about 50% larger than the wires in the thermocouple. Calculations 3 of Aeronautics and Astronautics ofthetimeconstants forjunctionballswithadiameter of 0.015 to 0.019 in. showed that these thermocouples have a time constant of about 1/2 sec. would Pressure/Flow Transducers. An absolute pressure transducer measured the pressure upstream of the thermal barrier while a di fferential transducer measured the pressure drop across a specimen. Flow through the thermal barrier was measured using a 0 to 100 SLPM flowmeter. Data was acquired from all of this instrumentation at a sampling rate of 10 Hz using Keithley data acquisition hardware and Labtech Notebook software. For each test, a 5-in. thermal barrier specimen was prepared and installed into the groove in the fixture. The 14 thermocouples that measured the surface temperature along the specimen were slipped into the outer sheath layer of the thermal barrier and adjusted so that they were spaced properly. To prevent parasitic leakage, the plenum chamber ()-ring was then positioned so that it was snug against the ends of the thermal barrier. The vacuum pump was turned on for several minutes, to cause the pressure drop and to achieve a steady flow rate through the specimen before applying the torch. The oxyacetylene torch was adjusted until a neutral flame was formed. The torch was slid along a machined groove until it was properly positioned in front of the hole in the iris plate. The torch was left on the specimen for 30 or 60 sec. and then pulled away from the fixture and shut off. Sometimes repeat tests were performed on the same specimen to examine the effects of repealed |'lame exposures. Torch nozzle spacing to the iris plate proved to be important in controlling the maximum hot side temperature without melting the center Type B thermocouple (platinum-rhodium, Tmclt = 3308 °F). Torch spacings for Carbon-3 and Carbon-6 were 0.265 and 0.160 in. respectively. Flow Tests Flow tests were perR_rmed on the thermal barriers in a high temperature flow and durability test rig shown schematically in Fig. 3. The test rig is capable of operating at temperatures from room temperature to 1500 °F, pressures between 0 and I(X) psig, and flows of 0 to 3.5 SCFM (standard cubic feet per minute, conversion I SCFM = 28.3 SLPM). Spccimcn length was 7.50-!-_0.05 in., and thc thermal barriers were mounted into a groove in the piston. The free ends of the specimens were joined together in the piston groove using a I/4 in. lap joint. Preload was applied to the specimens through a known interference fit between the thermal barrier and the cylinder inner diameter. To vary the amount ofpreload, the interference fit was modified by mounting different thicknesses of stainless steel shims behind the specimen in the piston groove. During flow testing, hot pressurized air entered at the base of the cylinder and flowed to the test specimen that sealed the annulus created by the cylinder and piston walls (0.007 in. American Institute radial gap). The durability of the thermal barriers at high temperatures was examined by subjecting them to scrub cycles in which the piston and thermal barrier were reciprocated in the cylinder. Flow data was recorded before scrubbing at temperatures of 70 and 500 °F and after scrubbing at 70, 500, and 900 °F. Specimens were subjected to ten scrub cycles at 500 °F. At each temperature, flow data was recorded at pressures of 2, 5, 10, 30, 60, 90, and 100 psid (or as high as could be recorded within the limits of the flowmeters) with the downstream pressure at ambient pressure. Primary and repeat flow tests were perlormed on the Carbon-3 and Carbon-4 designs for a diametral or linear compression of 0.040 and 0.050 in. (20 and 25% linear compression) and on the Carbon-6 design at linear compressions of 0.052 and 0.065 in. (20 and 25% linear compressions). A detailed description of the hardware and procedure used to perform these tests can be lbund in the papers by Steinetz et al. I and Steinetz and Adams. 2 Compression Tests Compression tests were performed to determine thermal barrier preload and resiliency behavior at room tcmpcrature using a precision linear slide compression test fixture shown schematically in Fig. 4. A 1 1/2-in. long specimen was loaded into a stationary grooved specimen holder, and an opposing plate was compressed against the specimen. Stainless steel shims were placed in the groove behind the specimens to vary the amount of linear compression. The amount of compressive load on the specimen was measured versus the amount of compression. Multiple load cycles were applied to the specimen belore the preload data point was recorded to remove effects of the hysteresis and permanent set that accumulate with load cycling of the specimens. Most permanent set occurred within the first tbur load cycles. A pressure sensitive film mounted on the opposing plate was used to determine the contact width of the specimen as it was compressively loaded. The footprint length (nominal I in.) and width at the end of the fourth load cycle were used along with the measured load versus compression data to calculate the estimated prcload and residual interference corresponding to a given linear crush value. I Residual interference is defined as the distance the specimen will spring back while maintaining a load of at least I Ib/in. of specimen. Compression tests were per|ormed on the Carbon-3 and Carbon-6 designs to determine the specimen preloads corresponding to the linear crushes used in the flow experiments. Tests were performed at compressions of 20, 25, and 30% of each specimen's overall diameter. Primary and repeat compression tests were performed. The hardware and procedure used to perform these tests are described in detail by Steinetz et al.I 4 of Aeronautics and Astronautics Subscale Rocket "Char" Motor Tests As part of the development process of the thermal barrier, Thiokol Corporation performed tests using a subscale (701 bm) rockct "char" motor. In these tests, the NASA Carbon-6 0.260-in. cross-sectional diameter thermal barrier impeded hot gas flow through an intentional circam ferential defect between rocket-case insulation blocks. The thermal barrier compression was 20%. The insulation blocks were modi fled to accommodate a 5 1/8-in. diameter thermal barrier. The 0.060-in. defect was much larger than any defects that would normally lorm through the gap-fill material in the actual rocket nozzle joint, but this size was chosen to force gas flow through the thermal barrier under very extreme conditions. Burning solid rocket propellant, the rocket fired for I I sec. and generated 900 psi pressures and 5000 °F (estimated) chamber temperatures. Hot gas l'lowed to the thermal barrier while upstream and downstream temperatures and pressures were recorded. The char motor incorporated an outboard plenum chamber, or reservoir, to simulate the volume (80 in. J ) between the thermal barrier and the Viton ()-ring seals. This reservoir ensured that flow would pass through the thermal harrier. The reservoir started at ambient pressure and then quickly reached chamber pressure, simulating the actual RSRM ,joint fill-time. After the volume between the thermal barrier and Viton ()-ring pressurizes in the rocket nozzle joint, charring risk to the Viton ()-ring is virtually eliminated. Results Burn Tcsl Results The amount of time to burn through each type of thermal barrier is shown in Fig. 5. In this figure, the number of specimens that were tested is given next to the name of each thermal barrier type, and the average burnthrough time is found above each bar. As shown previously by Stcinctz and Dunlap, 3 carbon fiber thermal barriers were the most burn-resistant. Figure 5 summarizes the earlier tests done on I/8-in. diameter stainless steel rods. Viton ()-rings. and all-ceramic braided rope seals. It also shows the burn times of the I/8-in. diameter (Carbon-1, Carbon-2. and Carbon-2A) and 0.200-in. diameter (Carbon-3 and Carbon-4) carbon thermal barriers as well as new data on the burn time of the 0.260-in. diameter Carbon-6 design. The I/g-in. diameter designs all endured the 55(X) °F oxyacetylene torch tor about 2 rain, Even more impressive burn times were seen h)r the 0.200-in. diameter designs at about 6-1/2 rain. This is more than three times the Shuttle solid rocket motor burn time of 2 min. 4 sec, However. an increase in diameter to 0.260 in. did not produce an increase in burn time. Carbon-6 at 0.260 in. in diameter had a similar burn time to the 0.2(X)-in. diameter designs at about 6-1/2 rain. Like the other carbon thermal barriers. Carbon-6 was soft and flexible after removal from the flame, even in the area and Discussion Thermal Barrier Porosity Measurements Measured values for thermal barrier density and porosity at 20_ compression are presented in Table II lbr the Carbon-3 and Carbon-6 thermal barrier designs. A 20% compression level was chosen, as this is the compression level selected lbr the nozzle-to-case ,joint thermal barrier. The densities/porosities of braided structures arc important for understanding their thermal and flow response characteristics. Carbon-3 had a higher density (0.041 Ib/cu.in.) and a lower porosity (0.37) than did Carbon-6 (0.032 Ib/cu.in. and 0.50, respectively). This can be attributed to the differences in braid architecture between these two designs as shown in Table I. Carbon-3 had a core composed of ten uniaxial 12K yarns of Grafi134-700 carbon fibers-a large fraction of its cross-section, while Carbon-6 only had one 12K yarn in its core. Carbon-6 had ten sheath layers of braided carbon fibers, while Carbon-3 only had five layers. Carbon-6 also had a lower sheath braid angle and fewer carriers per sheath layer to produce a softer, more flexible thermal barrier. Because the uniaxial fibers in the core pack together much better than the braided fibers that cross over each other in the sheath, the Carbon-3 design with a American greater percentage of core fibers is naturally more dense and less porous. Steinetz and Dunlap 3 showed previously that the density of a braided carbon thermal barrier was inversely related to the number of sheath layers. affected by the flame, with no evidence of charring or melting. All of the non-carbon specimens showed signs of charring or melting after removal from the flame, and many became very brittle in the area that was burned. The similarity in burn time between Carbon-6 and thc smaller-diameter Carbon-3 and Carbon-4 thermal barriers is believed to be related to the difference in porosity between these designs. As shown in Table II, Carbon-6 is more porous than Carbon-3 even in a compressed state. Steinctz and Dunlap 3 theorized that the mass-loss mechanism during the oxyacetylene torch tests was carbon oxidation. Depending on material type, carbon fibers begin to oxidize at temperatures in the range of 6(X) to 900 °F. 7"9 The oxyacetylene torch burning at 5500 °F is hot enough to cause oxidation to occur, but too cool for carbon sublimation that occurs at 6900 °F. 10 It is believed that the looser, more porous braid of Carbon-6 allowed more of the hot, oxidizing torch flame to pass through it. This allowed oxidation to occur more rapidly in the innermost fibers of Carbon-6 than in the less porous Carbon-3 design. Even though there were more carbon fibers in the larger Carbon-6 design, they were cut through more quickly because they were exposed sooner to hot, oxidizing gases. These results indicate that burn/oxidation 5 Institute of Aeronautics and Astronautics resistance isdependent onboththermal barrierdiameter andporosity. Products of combustion in thesolidrocketmotor include liquidalumina (A1203 jandgaseous CO,CIO2,CI, HCI.andH.,,noneof whichareoxidative. Hence, it is believed thattheneutralflameinambient air(oxidizing) isaconservative (i.e.,moreaggressive) environment for performing material screening burntests.It isexpected thatoxidation rateswithinthercx:ket environment willbe slowerthanthose exhibited herein. Temperature Drop Test Results Temperature drop tests were performed on the Carbon-3 and Carbon-6 thermal barrier designs using the test fixture described that measured the temperature drop across and along the thermal barrier in a compressed state when subjected to the flame ol'an oxyacetylene torch. Figure 6 shows temperature versus time traces for a test performed on a Carbon-3 specimen. Data recorded from the center thermocouple and the three thermocouples to the right of center on both the hot and cold sides of the specimen are presented. Data from the thermocouples to the left of the center thermocouple is not shown in this figure for clarity. In general, the left and right sides produced symmetric data. Also shown in the figure is the temperature trace from the "cold bulk" (Tbulk) thermocouple that measures the air temperature 1/4 in. downstream of the specimen. For sensitivity purposes, we moved the Tbulk thermocouple spatially to see if we were missing any local "'hot-streaks,'" and we did not find any. Figure 7 shows temperature traces for a test performed on a Carbon-6 specimen. Examining Figs. 6 and 7, it can be seen that the center thermocouple on the hot side (Tho I) and the center thermocouplc on the cold side (Tcold) of the thermal barrier each recorded the hottest temperatures on their respective sides. This is expected as these thermocouples are directly in line with the hottest part of the torch flame as it passes through the hole in the iris plate. These figures also show that the temperature got progressively cooler from the center thermocouple to the R l, R2, and R3 thermocouples on the hot and cold sides of the specimen. This was also expected as the temperature decayed with movement further away from the center heat source. Figures 6 and 7 show that there was a lag between increases in temperature on the hot and cold sides of the specimen. When the torch was applied to the thermal harrier, the hot side thermocouples instantly registered the increase in temperature. The insulating properties of the thermal barrier delayed heat conduction to the cold side, so the cold side thermocouples did not register an increase in temperature until several seconds after the torch was applied. The cold side temperatures measured were signiticantly lower than the hot side temperatures, as will be American Institute discussed below. Alter the torch was pulled away from the specimen, the hot side thermocouples instantly showed a decrease in temperature. The cold side thermocouples, though, continued to increase lor 3 to 5 sec before beginning to decrease in temperature. Comparing the hot side temperatures in Figs. 6 and 7, one notes fluctuations in temperature tor Carbon-6 but not Carbon-3. The origin of this fluctuation is unclear at this point, but we could find no system source of the variation (e.g. thermocouple integrity, etc.). Figure 8 shows the temperature drop across specimens of Carbon-3 and Carbon-6 for flame applications of-30 sec. The temperature drop was calculated as the difference between the temperature recorded by the hot side center thermocouple and the cold side bulk temperature (Tbulk). Over the 30-see. torch applicatiom the temperature drop across the Carbon-3 specimen dropped from a high of 2870 to 2680 °F by the end of the test. This drop was caused by a steady rise in the cold side bulk temperature while the hot side temperature remained nearly constant. Carbon-6 exhibited a temperature drop in the range of 2980 to 2600 °F. The uneven nature of the Carbon-6 trace is duc to fluctuations in the hot side temperature, as noted above. As shown by these figures, both Carbon-3 and Carbon-6 thermal barrier designs caused a comparable temperature drop across the thermal barrier over a 30-see. torch flame application. Figure 9 illustrates the symmetry of the temperature drop data [br Carbon-3 and Carbon-6. Figure 9(a) shows the temperatures recorded by the seven hot and cold side thermocouples that were in contact with the surface of a Carbon-3 specimen 15 sec. into the test. Though the downstream volume in the nozzle-to-case joint of the Shuttle solid rocket motors is expected to fill in <10 see., 15 scc. was chosen to include a safety factor of 5 sec. Figure 9(b) shows similar data for a test performed on Carbon-6. Both figures show the temperature distribution from left to right across the hot and cold sides of the thermal barriers. The center thermocouples on the hot and cold sides correspond to a position of zero. Thermocouples to the left of center have a negative position value, while those to the right have a positive value. Both figures show a temperature distribution that is close to symmetric around the center thermocouples. Figure 9(a) shows that the data lor this Carbon-3 test is shifted slightly to the right. Both figures show a temperature drop of about 2300 °F between the hot (Tho t) and cold (Tcokl) center thermocouples in contact with the surface of the specimens. Jet Spreading. The jet spreading capability of Carbon-3 and Carbon-6 is also shown in Fig. 9. Although the hot (3000+ °F) torch was focused into a narrow (0.084-in. diam.) column, the thermal barrier spread the heat at least I in. on either side of the center thermocouples. Figure 9(a) shows that for Carbon-3, temperatures I/4 in. away from the center hot side thermocouple were about 6 of Aeronautics and Astronautics 1200 °F on the left side and over 2000 °F on the right side. Hot side data for Carbon-6 in Fig. 9(b) show a similar trend with temperatures I/4 in. away from center over 2200 °F. Cold side data from both Figs. 9(a) and (b) show that the hot gas ,jet was reduced in temperature and diffused. Reducing the unit thermal energy per area is beneficial in preventing hot gas effects on the downstream ()-rings. Focused Jet Endurance Tests. Table Ili and Fig. 10 summarize the results of repeated temperature drop tests performed on single specimens of Carbon-3 and Carbon-6 to examine their endurance alter multiple applications of the oxyacetylene torch. For both thermal barrier designs, a single specimen was subjected to the torch flame for two 30-sec. periods followed by two 60-see, pericvds. The exposure times of 30 and 60 sec. are longer than the,joint cavity fill time of 10 sec. but were selected to examine the thermal barrier's insulation and flame resistance properties. After each exposure, the specimen was photographed (with fixture cover plate removed) to record any specimen damage before the next test was performed. For reference, the Carbon-6 specimen was also exposed to a 20-sec. flame application before these endurance tests, and no damage was observed. Table III shows several important temperature measurements for each test after 15 sec. as well as the flow through the specimen at fifteen seconds, the maximum bulk temperature reached during a test, and the amount of recession on the hot side of the specimen after the final flame exposure. The data for Carbon-3 shows that tests 30, 3 I, and 32 were almost identical. Each showed a maximum hot side temperature slightly above 3000 °F and a temperature drop (Thot - Tbulk) of ovcr 2800 °F. The only difference between these tests was the higher maximum bulk temperature of 500 °F in tcst 32. This was due to thc longer flame exposure time that allowed the bulk temperature to keep increasing tor 60 sec as compared to the 30-see exposures in tests 30 and 31. The maximum hot side temperature in test 33 only reached 2590 °F compared to 3(900+ "Fin the other tests. This caused lower temperaturc differences across the specimen and lower bulk temperatures. For all four tests, the highest bulk temperature after 15 sec. was 230 °F. This is well below Viton's short term maximum operating temperature limit of 600 °F. l I Even the maximum bulk temperature of 500 °F recorded after 60 sec. of flame exposure was within the limit. Figure IO(a) shows the hot side of the Carbon-3 specimen after all four flame exposures. No damage can be seen after the first three tests with little if any damage evident alter the final test. As shown in Table lit, there was a recession of 0.029 in. ( 13% of the compressed cross-section) measured alter 180 sec. of exposure. The thermal harrier should never experiencc such a prolonged exposurc to.jets of hot gas in the actual rocket application. American Institute The endurance tests performed on Carbon-6 revealed results slightly different than tor Carbon-3. After 15 see., the maximum temperature ranged from 2520 to 2730 °F with temperature drops (Thot-Tbulk) that ranged from 2240 to 2560 °F. The maximum bulk temperature after 15 see. was 280 °F, slightly higher than that for Carbon-3 but still well below the Viton ()-ring temperature limit. The Carbon-6 series revealed a slightly higher maximum overall bulk temperature of 620 °F that occurred in the final test after a 60-see. flame exposure. This temperature is about the maximum that the ()-rings can withstand for a short period of time, but as mentioned previously, the thermal barrier should not experience such a long flame exposure in the rocket. Figure 10(b) shows the hot side of the Carbon-6 specimen after all lour flame exposures. Very little damage can be seen after the first test. but the amount of damage to the specimen increased to a maximum recession of 0.092 in. (30% of the compressed cross-section) after the final test. This recession likely contributed to the increased maximum bulk temperature in the final test. These temperaturc drop tests were all performed in a more aggressive oxidizing environment than the thermal barrier would experience in the rocket. The amount of damage observed on thc Carbon-6 specimen after 2(X) sec. of flame exposure would not bc expected to _v,:cur in a less oxidizing environment with much shorter hot gas exposures. For both series of tests, the flow through the specimen was almost identical from test to test. Flow rates through Carbon-6 were higher than those through Carbon-3 as is expected since Carbon-6 is more porous than Carbon-3 (Table I1). Flow Test Results Flow rates (measured for Carbon-3, Carbon-4, using the piston flow rig, Fig. 3) and Carbon-6 at 20 and 25c)_ linear compression are summarizcd in Fig. I I at 60 psid and 70, 500, and 900 °F after scrubbing and 70 °F belorc scrubbing. Application of the thermal barrier in the Shuttle solid rocket motor nozzle-to-case joint involves predominantly static (e.g. no scrubbing) loads. As shown by the flow results, flow resistance increased with higher compression levels. Figure 11 shows that the flow rates for Carbon-6 were higher than those for Carbon-3 and Carbon-4 at 60 psid at each temperature and compression level. Carbon-6 flow rates were 2. I to 2.9 times higher than Carbon-3 flow rates and 1.7 to 2.3 times higher than Carbon-4 flow rates at comparable temperatures and compression levels. This difference is due to differences in braid architecture between these thermal barrier designs. The difference in flow rates between Carbon-3 and Carbon-4 was attributed to Carbon-4 incorporating larger core fibers resulting in higher seal porosity than Carbon-3. 3 Carbon-6 incorporating multiple sheath layers 7 of Aeronautics and Astronautics hasa higherporositythanCarbon-3 (Table1I)andis therefore more permeable. Discussions between theauthors androcketmanufacturer Thiokolhaveindicated thatthe thermal barriers have highenough permeability topermit thejoint-cavities tofill inacceptable times. Effect of Temperature. Figure II shows that flow rates dropped for each thermal barrier as the temperature was increased. This phenomenon is explained by the relationship that gas viscosity increases with temperature, ,IJo_ T 2/3. Thus, as the viscosity of the gas flowing through the thermal barriers increased, the flow rate decreased. 2 Effect of Hot Scrubbing. Thermal barrier flow rates typically rose after hot scrubbing during flow tests. Alter 500 °F testing Carbon-6 flow rates rose as much as 20_ as compared to the flow rates belore scrubbing. Post-scrub room temperature flows lot all thermal barriers were done after time spent at 500 °F (2 hr) and 900 °F ( 1.5 hr). Postscrub r_om temperature flow rates for Carbon-3 as much as doubled as compared to their pre-scrub values. Carbon-6 exhibited similar flow growth after scrubbing but tlows for pressure differentials of 60 psid were not within the range of the flow meter used. It is believed that much of the flow rate increase is due to oxidation that occurred while the specimen soaked at these high temperatures. No major visible damage due to scrubbing was observed on any of the thermal barrier designs at the conclusion of the flow tests. Only minor fraying was observed at the specimen ends in the lap joint. Temperature exposure tests performed on carbon fiber thermal barriers 3 showed that short lengths of carbon thermal barrier lost weight when heated in a furnace at different temperatures tbr two-hour exposures. This supported the theory that the carbon thermal barriers oxidized when exposed to temperatures of 9(X) °F for extended periods of time. and the associated weight-loss contributed to the increased flow rates after scrubbing. Compression Test Results Table IV summarizes the results of the compression tests performed on Carbon-3 and Carbon-6 and includes the measured contact width, preload, and residual interference Ior each amount of linear compression, or crush, tested. Contact Width. The contact width increased for the Carbon-3 and -6 designs as the amount of linear crush was increased. The thermal barriers continued to spread and flatten out as they experienced larger amounts of compression. In each test, the footprint pattern left on the pressure sensitive film after a compression cycle was solid and continuous. This indicates that during a flow test continuous contact is made between the walls of the flow fixture and the thermal the specimen. barrier, minimizing American leakage past Institute The contact width at each compression level for Carbon-6 was over twice as large as it was for Carbon-3 even though the diameter of Carbon-6 was only 1.3 times larger than tor Carbon-3. This shows that Carbon-6 had a softer, more compressible braid architecture than Carbon-3 allowing Carbon-6 to spread out more as it was compressed. Preload. The amount of preload or footprint contact pressure increased with the amount of linear crush. However, Carbon-6 had preloads that were 1/6th to 1/9th those ft_r Carbon-3 at each compression level. As a result, Carbon-6 will cause lighter loads on the adjoining rubber J-leg element. The reason for this difference in preload is believed to be related to the architectures of these thermal barrier designs (Table I). In Carbon-3 having a tightly packed core of uniaxial fibers, there is little room lor individual fibers to move with respect to one another when they are compressed. In contrast, in Carbon-6 the sheath fibers are oriented at an angle with each other and arc better able to slide past each other when the thermal barrier is compressed. Residual Interference. As with the contact width and preload, thermal barrier residual interference or spring back also increased as percent linear crush increased. Although contact width and preload were quite different for Carbon-6 and Carbon-3, residual interference scaled with diameter lor these two designs. Increasing thermal barrier diameter by a factor of 1.3 from 0.200 to 0.260 in. resulted in an increase in residual interference by that ratio for each level of compression. Residual interference for Carbon-6 was 0.025 in. even for the lowest compression (20%.) and meets the design requirement to Iollow nozzlc joint movement during Shuttle solid rocket motor operati on, as discussed with rocket manufacturer Thiokol. Comparison of Carbon-3 and Carbon-3 and Carbon-6 temperature drop comparison somewhat greater insulating showed less recession than Carbon-6: Other Factors both performed well in thc tests. Carbon-3 did offer effects than Carbon-6 and Carbon-6. We believe the higher density of Carbon-3 is an important reason for these results. However, there are many other factors to consider when deciding between these two braid architectures. Carbon-6 is braided using larger tows or yarns that permits faster and therefore most cost-effective production. Carbon-6 is a more flexible braid that makes it easier to spool for shipment and more accommodating during installation. The current tests combined with other planned rocket motor and joint-simulation tests will enable Thiokol and NASA to decide on the optimal braid architecture tot the thermal barrier. 8 of Aeronautics and Astronautics Results of Thiokol Char Motor Tests on Carbon Thermal Barrier Thiokol tested a 0.260-in. diameter Carbon-6 thermal barrier for NASA in a subscale rocket motor to verify that it would withstand the Shuttle solid rocket motor environment. The subscale motor, or "char" motor, simulates the effective slag barrier. The inset photo in the figure shows a close-up of an area where slag was trapped by the thermal barrier, preventing it from reaching the downstream ()-rings. Minor fraying occurred in the area immediately around the lap.joint during disassembly, but the specimen is otherwise in good condition. thermal conditions of the full-scale motor by burning solid rocket propellant at corresponding chamber pressure and temperature conditions. The thermal barrier was placed into an intentional gap defect between the phenolic insulation blocks, as shown in Fig. 12(a). The combination of an outboard plenum chamber and the 0.060-in. circumferential Comparison gap extending both upstream and downstream of the thermal barrier ensured thal hot gas flow would pass through the thermal barrier. Throughout the test duration of- 1 I see., a significant drop in temperature was measured across the thermal phenolic material to simulate the material and boundary conditions that the thermal barrier would be exposed to in these other configurations. The thermal barrier specimens were subjected to 209b compression as they were in the char motor test and as planned for the rocket. The flame of barrier. Figure 12(b) shows that the maximum temperature seen on the hot side of the thermal barrier was over the oxyacetylene torch that was used for the temperature drop tests was directed through a 0.084-in. diameter hole in an iris plate to simulate a hot gas jet that the barrier could be exposed to in the rocket. Flame exposure times were 32(X) °F, while the cold side temperature reached about 950 °F. Thus, a temperature drop of about 2200 °F occurred across the 0.260-in. diameter thermal barrier. Pressure readings upstream and downstream of the thermal barrier and in the reservoir confirmed that there was gas flow across the thermal barrier. The thermal barrier diffused the focused nature of the hot gas jet, further reducing the jet's potentially damaging effects on downstream Viton ()-rings in the actual Shuttle solid rocket unotor. Although the 950 °F temperature recorded downstream of the thermal barrier is still higher than the temperature limits of the Viton nozzle ()-rings, the char motor subjected the thermal barrier to more aggressive conditions than would ever occur in the actual Shuttle solid rocket motor, for the following reasons. First the gap defect was purposely oversized at 0.060 in. to force flow through the thermal barrier. In the actual nozzle ,joint, the gap between adjoining blocks of insulation would be narrower as the pieces of insulation are basically in contact with each other. The narrow gaps between the phenolic insulation would significantly cool the incoming gas temperature impinging on the thermal barrier and would therefore lower the temperature of the gas that reaches the Viton ()-rings. Furthermore, the downstream temperature in the char motor test was recorded immediately downstream of the thermal barrier. The ()-rings in the rocket nozzle .joint are located several inches further downstream of the thermal barrier, allowing additional heat to from the gas before reaching the ()-rings. Figure 13 shows the thermal barrier removed from the char motor. There was burning or charring of the thermal barrier. Fig. 13 shows that the thermal barrier also American be removed after it was no apparent In addition, acted as an Institute of Thiokol Char Motor Test Results to NASA Temperature Drop Test Results The fixture used to perform the temperature drop tests on the Carbon-3 and Carbon-6 thermal barriers was modelled after the char motor and the shuttle nozzle-tocase joint thermal conditions. The fixture was made out of intentionally longer than they would be in the rocket application to simulate extreme heating conditions. Considering the results of Fig. I0 (NASA temperature drop fixture), tests were performed with hot side temperatures ranging from 25(R) to nearly 32(X) °F. Carbon-6 temperature drops ranged from 2240 to 2560 °F-I 5 sec. into the test. These were somewhat greater than the 2200 °F temperature drop exhibiled by Carbon-6 in the char motor. The main reason for this difference is that 9(X) psi pressures were generated by the char motor, while only 10 psid pressures were applied across the thermal barrier in the temperature drop tests. The higher-pressure char motor test caused more hot gas to tlow through the thermal thereby raising the downstream temperature causing a smaller temperature drop. Though there are some differences in the absolute results, the authors believe the laboratory temperature-drop test fixture simulates many of the key factors at work in the rocket. The laboratory setup permits quick and easy comparisons between competing architectures and can be used to generate thermal data to anchor thermal correlations under development. Summary and Conclusions The 55(X)+ °F combustion gases in the Space Shuttle solid rocket nnotor are kept a safe distance away from the assembly .joint seals by thick layers of insulation and by special compounds that fill the joint split-lines in the insulation. The current nozzle-to-case joint design incorporates primary, secondary and wiper(innermost) ()-rings 9 of Aeronautics and Astronautics and polysulfidejoint-fill compound. In the current design, I 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 (successfully used in the field and igniter joints) and the thermal barrier Carbon-6 described herein. The thermal resistance of two NASA thermal barriers, denoted Carbon-3 and Carbon-6. was assessed Ik_rmed to measure the temperature drop across and along the thermal barriers in a compressed state when subjected to the flame ofan oxyacetylene torch. Flow and durability tests were conducted on the thermal barriers to examine their leakage characteristics and durability at ambient and high temperatures. Room temperature compression tests were pertormed to determine load versus linear compression, preload, contact area. and residual interference/ resiliency characteristics. Subscale rocket "'char" motor tests were performed in which hot combustion gases were directed at the Carbon-6 thermal barrier to assess its thermal resistance in a rocket environment. The current tests with other planned rocket motor and joint tests will enable Thiokol and NASA to decide on the optimal braid architecture for the thermal barrier. Based on the results of the current conclusions are made: tests, the following I. The Carbon-6 (0.260-in. diam.) and Carbon-3 (0.20-in. diam.) thermal barrier resisted the 5500 °F flame of an oxyacetylene torch for over 6 min before burn through, greater than three times the Shuttle solid rocket motor burn time. 2. Carbon-3 and Carbon-6 thermal barriers were excellent insulators causing temperature drops through their diameter from 25(X) to 2800 °F, depending on test parameters. Gas temperature I/4" downstream of the thermal barrier were within the downstream Viton ()-ring temperature limit of <600 °F. 3. The Carbon-6 thermal barrier design performed extremely well in subscale rocket "char" motor tests that subjected it to hot gas at 3200 °F for an I I-see. rocket firing, simulating the maximum downstream joint-cavity filltime. The thermal barrier reduced the incoming hot gas temperature by 2200 °F in an intentionally oversized gap American References by exposing them to an oxyacetylene torch at 5500 °F and measuring time for burn through. Temperature drop tests were per- combined simulation defect, spread the incoming jet flow, and blocked hot slag, thereby offering protection to the downstream O-rings. 4. Laboratory burn, temperature drop, flow, and compression tests and subscale rocket "char" motor tests demonstrate the thermal barrier's feasibility for use in rocket applications and qualify it tbr comprehensive motor evaluation. Institute ISteinetz, B.M.,Adams,M.L., Bartolotta, P.A., Darolia, R., and ()lsen, A., "High Temperature Braided Rope Seals tbr Static Sealing Applications," NASA TM- 107233, rev., July 1996. 2Steinetz, B.M.. and Adams, M.L., "Effects of Compression, Staging, and Braid Angle on Braided Rope Seal Performance," NASA TM-107504, July 1997. 3Steinetz, B.M.. and Dunlap, P.H.."Feasibility Assessment of Thermal Barrier Seals for Extreme Transient Temperatures," NASA TM-208484, July 1998. 4Thiokol report TWR-7319 I. "RSRM-45A Nozzle Joint No. 3 ()-ring Erosion Investigation Team-Final Report," October 28, 1996. 5Rogers, W.P.. "Report of the Presidential Commission on the Space Shuttle Challenger 1986. 6Ballis. W., ASM Handbook, Accident,"Junc Volume 6: Welding, 6, Brazing, and Soldering, ASM International, 1993, pp. 281-290. 7Bahl, O.P. and Dhami, T.L., "Oxidation Resistance of Carbon Fibers," High Temperatures - High Pressures. Vol. 19, pp. 211-214, 1987. 8Eckstein, B.H. and Barr, J.B., "An Accelerated Oxidation Test for Oxidation Resistant Carbon Fibers,"MaterialsProcesses: Twentieth The Intercept International Point; Proceedings of the SAMPE Technical Conference, Minneapolis, MN, Sept. 27-29, 1988. Covina, CA, Society for the Advancement of Materials and Process Engineering, 1988, pp. 379-391. 9Eckstein, B.H., "'The Weight Loss of Carbon Fibers in Circulating Air," 18th International SAMPE Technical Conference, October 7-9, 1986, pp. 149-160. l°Lide and Kehiaian CRC Handbook of Thermophysical and Thermochemical Data, CRC Press, 1994. pp. 25-31. I Iparker O-t4ng Handbook, 10 of Aeronautics and Astronautics Cleveland, OH, 1992. Barrier type Carbon- TABLE Core Size Diameter, in. I (I. 125 Carbon-2A 0.125 Carbon-3 0.200 Carbon-4 0.194 Carbon-6 0.260 Grafil h 34-700 12K Grafil 34-7011 12 K 34-711tl 3K Grafil 34-71111 12K 34-700 3K Grafil 34-7011 12K Amoco" P25 2K Grafil 34-71)tl 12K 2.76xl0 72110 1800 2.76xl0"* _ 72110 2.76x I1) "_ 4.4x 10 "_ 2.76xl0 Thermal _Porosity 2.8x1(1 _ 9 8 I 45 l0 Thornel T-3111) I K Thornel T-30() I K Thornel T-30() 1K T-300 3K All-Ceramic 61tl) 2.8x10 5 I 61111 2.8x I 0 _ 5 600 181111 2.gxl0 10 12 in I-2 24 in 3-5 12 in I-2 24 in 3-5 8 in I-5 12 in6-7 16 in 8-I1) 65 in I _' 61) in 5'" 65 in I _' 61) in 5 _" 17 in I "_ 45 in 2-111 of Diameter. in. 0.2011 0.260 I Exposure 31 32 33 Tesl number 35 36 37 38 30 [ 311 60 60 [ Exposure Per test, SCC 31) 311 60 60 5 8 I 45 600 2.8xl0 II) 8 I 45 "_ "_ a I 700 I 3.2xl0_ [ Thermal POROSITY barrier Carbon density,. Ib/cu.in. 0.1141 0.032 Ill.--TEMPERATURE time T,,,,al 15see. IP.,e,,.I A.'o,,,uUed. °E SCC sec 31) 2.8x111 a I 2 8 I I I ' AT fiber Porosity" densit?', Ib/cu.in 0.06,4 t).1164 0.37 0.50 = I - Pu_/Pcl . (a) Carbon-3 Test number 6(XI THERMAL BARRIER COMPRESSION layers TABLE DROP Temperature RESULTS Drop Test Results J Th,,,_ at T,.,, - Th,,it 15see, J at 15see, T,,,,_- T...... I at 15 sec, °F I I TEST °F Fiow at 15 sec. SCFM/m. I Th,,ik lnaxinlunl, I I Recession lest in. °F 31) 3070 I 2111 I 286t) 23311 111.14 310 611 121) 8( 3050 3020 2590 [ 2311 21X) 50 [ I 28211 28211 24411 23/_1 I 22511 .14 .H 3411 l- -I 500 ...... .14 340 time Accuruulated. sec 50 80 1411 200 921 Test Results (b) Carbon-6 Temperature 1)1 , - T,, u Flov, at TI,,,tat Ti,utk at Tt>, - Tt,_,tl, at'k,,, 15 sec. 15 sec, al 15 sec, at 15 sec. 15 sec. _'F SCFM/in. °F °F °F 2730 2690 25211 2701/ American angle. Braid degrees 6011 I 5 10 I Thornel T-300 IK II.--MEASURED Number sheath yarns per Numberof bund c I I 20% barrier t vpc Carbon-3 Carbon-6 carriers NumberofJ per ayer I I 21 4 layers Carbon Thornel' T-300 I K Thornel T-300 I K NTW.a C-2 J 11.1211 J NX 551F I 7°° I 3.2×10" I lilt, [ NX550 -'lxl0 in.=25gm. hGrafil type 34-700 carbon libers, Gralil Inc. product. 12K-12.0()0 tiber ends. _Thornel T-300 carbon fibers, Amoco Perlbrmance Products, Inc. producl. "Amoco P25 pitch fibers. Amoco Peffurmance Products, Inc. product. _NX 5511 = Nextel 550 fiber, 3M product, 73r;bAlzO, 27c/f SiO, TABLE MATRIX diameter, in? 4 720() 2.76x111 _ 181111 2900 CONSTRUCTION of yarns 7200 72011 BARRIER Sheath tliameter, in/ 0.114 Carbon-2 I.--THERMAL 171) 191) 2811 280 Institute 25611 25/)11 2240 2420 II of Aeronautics 2050 1960 17611 1700 11.24 0.24 0,25 11.24 and Astronautics Tt,,,,_ lllaxinlulll, °F 320 350 481) 621) I-- 01129 after Percent I-- --3 Recessiun after test in. Percent ...... ...... .... I).1)92 31) I s6 TABLE IV.--THERMAL BARRIER CONTACTWIDTH, PRELOAD, AND RESIDUAL INTERFERENCE FOR SEVERAL LINEAR CRUSH CONDITIONS Diameter, Nominal percent Linear Number of Contact Preload, Residual in. linearcrush, crush, sheath width, psi interference,_' percent in. lancers in. in. 0.2 20 0.040 5 0.063 310 0.019 25 .050 .082 490 .027 30 .060 .099 930 .I)33 Carbon-6 0.26 20 0.052 10 O.157 56 0.025 25 .(K_5 .192 81) ./)36 30 .078 .196 97 .041 "Residualinterferenceis defined as the distance that the thennal barrier will spring back while maintaininga load of al leasl 1 Ib/m of specimen. Themlal barrier type Carbon-3 (a) r Thermal barrier I _Vent port / /-- Leak check port / Secondary O-ring L Primary O-ring L.Wiper O-ring A Rocket centerline T-Throat (b) _assemblyr(_) .- Beadng assembly Exhaust flow _ ,- Forward exit ,/ cone assembly '//--_ / L_Nozzle inlet assembly / L Cowl assembly Section A - A / / /- Aft exit cone assembly _ Nozzle-tocase joint Figure 1.--Potential Shuttle solid rocket motor joint locations for thermal barrier. (a) Enlarged view of nozzle-to-case joint showing J-leg, wiper, primary, and secondary O-rings, leak-check port, and proposed thermal barrier location. (b) Overall nozzle cross-section (half view). American Institute 12 of Aeronautics and Astronautics Top cover plate (phenolic) removed for clarity _ r- Plenum chamber / _ F- Test _ _- Iris plate specimen / / _ / \ Vacuum pump roughing _• draws flow through flow meter/fixture Air _ _ "<'_ --- _,/_--_ _ //" - _-__ /_ ,_"_-_ Downstream T/C's Radiant stroke heating ITI 0.25 in. --" ,,, _/// --8 each, i .-/ Piston _T'/_ll --. " K Upstream ............. ambient Differential ........ 10-11 psid type K -_ of temperature drop Carriage _/ test fixture. nperature Force I liFT7]_ / Digital Specimen f" in piston ( groove _, "_ Y/A J _ g/) d_ , Square i/ 1 I_ _- Pressure sensitive radii -_ _ \_ _ 4-C_ _ _. r----m',_-r-m- - Specimen , Lap joint-7 holders_'l" 3. - ..( Load ceil (2) _|, .'r t L Ii. _,/, , _,/. _ /| "--" i'' "_ - -- Test specimen ' _Stationary plate , I T 0-100 plate) film I Insulation stationary grooves with corner . % indicator I( c°ntacts \ -_- plate 7Moving __- dia.lf-_ _mlZSin. torch _ :reesa22::pe *II-:- , oxyacetylene [_ IT I ":-Cylinder I die hole) S.S. Iris Figure 2.--Schematic :i , --- "-._ -_ i C 0 nng--' _ Miniature / Y-_/ \\_ surface _ (0.084 % z_Pressure / psi Hot air supply Figure 3._Schematic "/ of flow , /j , Figure 4._Schematic fixture. 13 American Institute , of Aeronautics and Astronautics / / / of compression / fixture. 400 387 399 350 ¢_ Diameters: .¢E ¢_ _ 300 C-3, C-4 = 0.2"; C-6 = 0.26" O ¢_ _ 250 E >_ Reference: o 200 O 1/8" nominal except 2 min 4 sec Shuttle solid _ e- rocket motor burn time -7 150 _: o_ ..... _'_ f>_"__ 100 ||,,1 ........ 0 7_ __'0 ,oo I I i i e" _ ¢n ¢n Figure 11 7 j"__133 < • _,o 5._Oxyacetylene 3500 y thermal Carbon torch burn test results ¢?, (n = number I barriers of tests performed). Distance from center (in.) 1 - Hot Center, Tho t 2 Hot R1 0.0 3 Hot R2 0.50 4 Hot R3 1.00 I1 30OO 5 Cold Center, 6 - Cold R1 2000 _ 1500 E _, Channel I 2500 _ 0.25 0.0 Tcold 0.25 7 - Cold R2 0.50 8 - Cold Bulk, Tbulk 9 - Cold R3 0.0 1.0 1000 500 0 0 10 20 30 Time 40 50 60 Figure 6._Temperature rise vs. time for simulated hot gas exposure upstream (hot) and downstream (cold) temperatures for Carbon-3. temperatures removed for clarity. American Inslitute 70 (sec) 14 of Aeronautics and Astronautics showing Left hand Channel 35O0 I I 2500 1 2 3 4 2000 5 6 7 1500 8 9 3000 E Distance from center (in.) 0.0 - Hot Center, Thot Hot R1 0.25 Hot R2 0.50 Hot R3 1.00 0.0 .....Cold Center, Tcold - Cold R1 0.25 - Cold R2 0.50 0.0 - Cold Bulk, Tbulk • Cold R3 1.0 1000 500 __ • :f ,_ °J-_7 ___J. 8_ 0 i 5 _--,_----__ ........... _ I I ! 01 ............ t 10 15 20 25 I 30 ..... _ I 35 L 40 I 45 _ 50 55 Time (sec) Figure 7.--Temperature rise vs. time for simulated hot gas exposure showing upstream (hot) and downstream (cold) temperatures for Carbon-6. Left hand temperatures removed for clarity. 3OO0 U. o F- Carbon-3 -1 2000 2500 6_ o t- / // Carbon- h_ o / 1500 "O .9,=1000 Q. E t-- 500 00 5I _ 1- I 15 I 20 I 25 30 - 35 40 45 Time (see) Figure 8.--Temperature drop vs. time Carbon-3 and Carbon-6. American Institute (Thot-Tbulk) 15 of Aeronautics at flame location for and Astronautics 5O 3500 3500 Hot side 3000 _" -_ Cold side (a) /_ 2500- 2500 .= "¢ 2000 2000 -_ Cold_ 0_ 500 0.084" _- (b) Hot side 3000 ¢z 1500 E 1000 ).084" / 1000 et diamet_...._. 5OO 500 0 -1.00-0.75-0.5-0.25 0 Position Figure 9.--Hot diameter) 0.25 0.5 0.75 I -0.75 b for thermal '-_1 0 Position barriers versus axial position (a) Carbon-3; (b) Carbon-6. at 15 seconds 16 American I -0.5 -0.25 (in.) side and cold side temperatures spreading 0 -1.00 1.00 Institute of Aeronautics and Astronautics 0.25 0.5 0.75 (in.) showing jet (0.082 in. 1.00 Test Test . rThermal !J, _ _ _ i barder_ Cold . _Thermal barrier 31 _ 36 Hot Cold . _Thermal barrier _ 32 37 Hot ,_ _ "tp_-,_'tll_ll_.j_ • _:_ : .... _ . ..... . _- Thermal 33 __ : 38 .ot (a) Carbon-3 ITest# Exposure Temperature Time Per test Accumulated 30 (sec) 30 (sec) 30 Drop Test Tbulk Tho t @15sec @15sec (°F) 210 (°F) 3070 (b) Carbon-6 Results Thot-Tbulk Test# @ 15sec Exposure Temperature Time Per test Accumulated (°F) 2860 35 Drop Test Results Thot Tbulk !Thot-Tbulk @ 15sec @15sec (sec) (sec) (°F) (°F) @ 15 sec (°F) 30 50 2730 170 2560 2500 31 30 60 3050 230 2820 36 30 80 2690 190 32 60 120 3020 200 2820 37 60 140 2520 280 2240 33 60 180 2590 150 2440 38 60 200 2700 280 2420 versus accumulated Figure lO._Thermal barrier condition and key temperatures 17 American lnstilule of Aeronautics and Astronaulics time. (a) Carbon-3, (b) Carbon-6. 0.70 0.66 0.60 [] 70 °F - Before scrub • 70 °F - After scrub [] 500 °F - After scrub [] 900 °F - After scrub 0.50 0.45 0.44 .c 0.40 o_ ,T 0.30 0.39 off 0.27 0.26 0. 0.20 .21 ;//ll 0.20 i 0.10 7"/ll __ • 0.12 0.11 -- -- -- ° °01JI --- 0.21 _-- 0.00 Percent compression[ 20% 25% Carbon-3 Figure 11 .--The &P = 60 psid. effect (0.20" dia.) of temperature, "Char" moto_ nozzle 20% test article / / Intentional / / Rocket Carbon-4 thermal barrier type, /-- "/'cold " /-- Tho t Carbon-6 scrubbing (0.26" dia.) and compression on flow, - 4000 / og- // g 3000 7"hot /__- _- 2000 flow °F ="-1 /-........ 06_,_ [ barrier - (a) Figure 25% J (0.20" dia.) 1000 Thermal 20% J [ _', (5000 25% i [ • ..... Tcol d / I 20 J tNI 11 -sec __J burn time-' I 30 Time, sec (b) 12.--Subscale configuration: Temperature (70 Ibm) "char" Carbon-6 data: thermal Upstream motor tests examining barrier impedes thermal hot gas flow (Thot) and downstream barrier (Carbon-6) through (Tcold) sides intentional of thermal barrier. 18 American Institute of Aeronautics and Astronautics effectiveness. joint defect (Courtesy (0.06 (a) Test in. gap). of Thiokol (b)' Corp.) RTV remaining from test --_. \ \ 5 1/8 in. Diam. Hot slag blocked by barrier Figure 13.--Photograph effectively Thiokol blocks of char motor 3200 thermal barrier (Carbon-6) °F gas for 11 sec. (joint fill time) after test. Thermal and blocks Corp.) 19 American Institute of Aeronautics barrier hot slag. (Courtesy and Astronautics of REPORT DOCUMENTATION PAGE FormApproved OMB No. 0704-0188 Public reporting burden for this collection of information is estimated to average 1 hour per response, includingthe time for reviewing instructions, searching existing data sources, galhering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this colleclion of information, includingsuggestionsfor reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson Davis Highway, Suite 1204, Arlington. VA 22202-4302, and to the Office of Management and Budget, Paperwork ReductionProject (0704-0188), Washington, DC 20503 1. AGENCY USE ONLY (Leave 12. REPORT blank) DATE 4. TITLE 3. REPORT TYPE AND DATES Technical 1999 June AND SUBTITLE Development !5. of Thermal Barriers lor Solid Rocket Motor Nozzle COVERED Memorandum FUNDING NUMBERS Joints WU-523-53-13-00 6. AUTHOR(S) Bruce M. Steinctz 7. PERFORMING Patrick ORGANIZATION National John and Acn_nautics H. Glenn Cleveland, and Space Aeronautics Washington, Jr. AND ADDRESS(ES) 8. PERFORMING ORGANIZATION REPORT NUMBER Administration Center at Lewis Field AGENCY NAME(S) E-11738 44135-3191 9. SPONSORING/MONITORING National Dunlap, NAME(S) Research Ohio H. and DC Space AND ADDRESS(ES) 10. SPONSORING/MONITORING AGENCY REPORTNUMBER Administration NASA 20546-(X)01 TM--1999-209278 AIAA-99-2823 11. 12a. SUPPLEMENTARY Prepared for Angeles, California, the 351h Joint June Subject This Propulsion 20-24, DISTRIBUTION/AVAILABILITY Unclassified 13. NOTES Conference & Exhibit cosponsored by AIAA ASME. SAE, and ASEE, Los 1999. STATEMENT 12b. DISTRIBUTION CODE - Unlimited Category: publication 37 is available (Maximum ABSTRACT Tbe Space Distribution: Shuttle from 200 solid the NASA Center for AeroSpace Nonstandard Information. (301) 621-0390 words) rocket motor case assembly joints arc sealed using conventional ()-ring seals. The 5500+°F combustion gases arc kept a salt: dis- lance away, from the ,,.cals by thick layers of insulation. Special joint-fill compounds are used to fill the joints in the insulaion to prevent a direcl llowpath Io the seals, ()n a number of occasions, NASA has observed in several of the rocket nozzle assembly joints hot gas penelmtion through defecfs in the jointfill compound. The current nozzle-to-case .joint design incorporates primary, secondary and wiper (inner-rnost) O-rings and polysulfidc joinl-lill cornlumnd, In the current design, Iou! to the wiper ()-ring results nozzle-to-case ,joint design of 7 motors experience hot gas to the wiper in extensive reviews bcfi)re resunfing flight. by irnplenventing a more reliable J-leg design ()-ring. Though the condition does not threaten motor sali_ty, evidence of hot gas NASA and solid rocket motor nmnufacturer Thiokol are working to improve the and a thennal barrier. This paper presents burn-resistance, temperature drop. flow, and resiliency test resuhs for several types of NASA braided c_on-fiber thermal harriers. Burn tests were perforn_cd to determine the lime to burn through each of the thermal harriers when exposed to the llarne of an oxy-acetylene torch (55(X) °F), representative of the 55(X) °F solid rocket motor combustion temperatures. Thermal barriers braided out of carbon fibers endured the flame for over 6 rain, three times longer than solid rocket motor burn time. Tests were perlc,17ned on two thermal barrier braid architectures, denoted Carbon-3 and Carbon-6, to measure the temperature drop across and along Ihe barrier in a compressed stale when subjected to the flame of an oxyacetylene torch. Carbon-3 and Carbon-6 thermal barriers were excellent insulators causing temperature drops through their dian,,eter of up to a 28OI) and 2560'F. respectively. Gas lenlpemlure 1/4" downstream of the thermal barrier were within the dov, nsb'eam Viton ()-ring temperature limit of 6(X) °F. Carbon-6 pertorrned extremely well in subscale rocket "char" motor tests when subjected to hot gas at 32(X) °F fi,',ran I I -see. rocket firing, simulating Ihe maxinmrn downstream joint cavity fill time, The thermal barrier reduced the incoming hot gas temperature by 2200 °F in an intentionally oversized gap delitct, spread the incoming jet flow. and blocked hot slag, thereby offering protection to the downstream ()-rings. 14. SUBJECT Seals: TERMS Space Carbon: Shuttle: 15. NUMBER Solid rocket motor; Fluid tlow, Design thermal barrier: OF PAGES 25 Test: Braid 16. PRICE CODE 20. LIMITATION A03 17. SECURITY CLASSIFICATION OF REPORT 18. SECURITY CLASSIFICATION OF THIS PAGE 19. SECURITY CLASSIFICATION OF ABSTRACT Unclassified Unclassified Unclassified NSN 7540-01-280-5500 Standard Prescribed 298-102 OF ABSTRACT Form 298 (Rev. 2-89) by ANSI sir. Z39-18