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Phase II of the S-6 test consisted of two solid propellant fire tests. Both tests were single proximity with a 0. 914-m (3-ft) cube of UTP-3001 solid propellant inhibited on five sides. In both tests, the propellant cube was burned on a sand pad with the uninhibited surface vertical and the test specimens placed within a few centimeters of this burning surface. Steel was introduced into both tests via a 9. Miim(3/8-in.) thick steel plate placed on top of the propellant cube prir r to ignition. The test specimen for the first test was an FSA with an Ir PICS and a thoria simulant fuel sphere. This specimen was preheated to approximately 1040 °C, placed next to the propellant surface and the propellant ignited. Both the GIS and thr? PICS survived the fire with little or no damage. The only effect on the GIS was a slight erosion of the graphite fibers by the steel. The damage '.o the PICS consisted mainly of a!umina and silica deposits on the surface. Neither the GIS nor the PICS was breached and no thoria was released. The test specimen for the second burn was a mockup of a damaged HSA, consisting of a simulant filled 8-paek inside pieces of the graphite aeroshell encased in tha remains of the Ir outer clad from the Phase I HSA impact test. Only four of the PISA'S had Ir PICS, the other four were unclad thoria spheres. The specimen was preheated to a PICS temperature of approximately 980 "C (1040 °C oven temperature), placed next to the uninhibited surface of the propellant and the propellaiit ignited. Though part of the Ir outer clad was consumed,, all of the GIS's and Ir PICS were intact and no thoria was released from the Ir clad PISA. 1.0 INTRODUCTION ponents to the sequential environments characteristic In preparation for the Lincoln Laboratory's of a catastrophic launch pad accident to evaluate LES 8/9 space mission, a series of ti-ats was pertheir capability to contain the PuO2 fuel. This formed to evaluate the nuclear safety capability of sequence of environments was to have consisted of the Multi-Hundred Watt (MHW) Radioisotope Therthe blast overpressure and fragments, followed by moelectric Generator (RTG) to be used to supply powthe fireball, low velocity impact on the launch pad, er for the satellite. One such safety test is Test No. and solid propellant fire. The blast overpressure S-6, Launch Pad Abort Sequential Test. The objecand fragments were subsequently eliminated from tive of this test was to subject the RTG and its comthis sequence. This report describes the procedures and presents the results of Phase II of Test S-6, Solid Propellant Fire. In this phase of the test, a simulant Fuel Sphere Assembly (FSA) and a mockup of a damaged Heat Source Assembly (HSA) were subjected to single proximity solid propellant fires of approximately 10-min duration. Steel was introduced into both tests to simulate the effects of launch pad debris and the solid rocket motor (SRM) casing that might be present in the fire zone. The rationale for using these two test specimens is based on the results of Phase I of Test S-6, Low Velocity HSA Impact. 1 In that test, a simulant HSA was impacted on a concrete pad at approximately 30 m/s (=*100fps). As a result of that impact, the HSA was badly damaged.with two FSA's being released. However, all of the Graphite Impact Shells (GIS's) were intact. It was, therefore, concluded that the most likely contact with burning solid propellant would be by an intact FSA rolling away from the HSA impact or the HSA itself impacting next to a piece of burning propellant. The fact that the HSA impact was almost inelastic would indicate that it is highly unlikely that either the HSA or one of the FSA's would ever end up on top of a piece of burning propellant. The small amount of dispersion observed in Phase I would tend to suggest that the double proximity is also very unlikely. Therefore, it was decided that the fire tests would be of the single proximity type with a 0. 914-m (3-ft) cube of solid propellant, simulating the full web thickness of a SRM, and involve an intact simulant FSA and a damaged HSA. 2. 0 TEST DESCRIPTION 2.1 Propellant The solid propellant used was UTP-3001 which is u&ed in the SRM boosters of the Titan III C vehicle on which the LES 8/9 spacecraft will be launched. UTP-3001 contains powdered aluminum and ammonium perchlorate dispersed in a rubber binder. Several other additives are present in small percentages. The propellant blocks were inhibited so as to burn only on one face to give the maximum burning time. The 0. 9m (3-ft) dimension corresponds to the web thickness of the Titan III C SRM's. The rationale is that a piece of an actual motor could not burn any longer than the time nee essary to burn through a full web thickness. The steel plate was added to the test because the actual SRM's havea steel casing and it is thought by some that a rocket abort or destruct could not produce a large piece of solid propellant without having steel 2 casing attached. In a previous fire test it appeared that steel introduced into the fire might have been the cause of observed fuel containment failure. 2.2 FSA Fire Test 2. 2.1 Test Specimen. The FSA used in this fire test was MHT-S5, made up of a thoria (J6% TD) simulant fuel sphere, a type WC Ir Post Impact Containment Shell (PICS), and a flight quality Graphite Impact Shell (GIS). This was one of the FSA's that was used to make up the "bottom" 8-pack of the HSA used in the Phase I impact test. A photograph of the Post Impact Sphere Assembly (PISA) from this FSA, taken after th. Phase I impact but before the fire test, can be seen in Fig. 2.1. The only damage suffered by this PISA during the Phase I test is the dent shown in the photograph. This PISA was placed in a GIS and preheated to approximately 1040°C prior to being placed next to the uninhibited uurface of the solid propellant. A second FSA, MHT-43, was equipped with a thermocouple to measure PICS temperature and placed near the burning surface of the propellant prior to ignition. This FSA was not preheated and was used only as a part of the temperature measuring device. 2. 2. 2 Test Setup. The FSA fire test was Bt.t up at the LASL K-Siie facility. A sand pad, approximately 3-m (9-ft) in diameter and about 0. 3-m (1-ft) thick was constructed in an open area just to the east of the drop tower used in the Phase I test. riiio.!Mu ct-rt) iiiiu<>r HTI>-:<OOI M«ud was |>lai'i<l iiciic the renter nF the san'J pad with tine inhibited surface IHWII, SH that the uninhibited surface was vertical and facing the niirthwust. A piece of steel plate, !1. !t<nmi(3/H-in.) thick, waiplaced on top of the piMpcUunt block with a few centimeters of the plate protruding over the uninhibited surface. A tube furnace, equipped with a remotely operated trap door, was suspended on a trip d, so that the PSA could be preheated and then placed near the uninhibited surface of the propellant. A ramp, made of wood, was used to get the heated KSA from the furnace to a position near tht< propeliant surface. The complete test setup can be seen in Kig. 2. '£. Also shown in the figure are the relative positions of the two '••'SA's just prior to propciiant ignition. The KSA in the foreground is the |>reheate<l test specimen. MHT-.IS, and the other is the non-preheated instrumented KSA, MHT-43. Just prior to ignition, the tube furnace and ramp were to have beer pulled out of the fire /.one by a cable attached to tlie lift at the drop tower (not shown in Kig. '£. 2). One a the furnac>' and ramp were clear, the propellant was to have been ignited '>y smokeless powder, keiosene-soaked excelsior, and electrically fired squibs placed at the base of the uninhibited surface. The !'"SA, MHT-43. was equipped witlia tungsten-rhenium thermocouple with the temperature measuring bead against the PICS. This KSA was placed near the base of the propellant, just prior to ignition with the thermocouple bead on the side of the PICS away from the fire, but was not prehea! ed. The thermocouple leads were passed through J. small hole in the GIS and under the sand of the pad. Away from the pad, the leads were covered with enough sand to insure that they would not be burned during the test. The thermocouple was connected lo a recorder and its output v.as monitored throughout

Safety test No. S-6, launch pad abort sequential test Phase II: solid propellant fire