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Reported study of composite-propellant burning summarizes recent advances in understanding behavior of propellant formulations based on ammonium perchlorate (AOP), binder, and aluminum in various proportions and particle size distributions. Approach presented incorporates adapted version of earlier model for monopropellant AP. Objective is to predict burning-rate characteristics of composite propellants at high pressure.
Two types of low vulnerability propellants are studied which are distinguished by whether the binder is a rubber, such as polyurethane or CTBN, or a plasticizable polymer such as ethyl cellulose or cellulose acetate. The former propellants are made by a partial cure extrusion process while the latter are made by the conventional solvent process. Emphasis is given to a cellulose binder (plasticizer) RDX composition. The type of binder used, the particle size of the RDX and the presence of small quantities of nitrocellulose in the solvent processed compositions have important influences on the mechanical and combustion characteristics of the propellant. The low temperature combustion is of particular concern because of potential breakup of the grains that can lead to instability.
It is reported that the neccessary technology has been developed and demonstrated for the manufacture of heat-sterilizable solid propellants which meet specific ballistic goals. It is shown that: (1) phosphate doping of ammonium perchlorate significantly enhances the thermal stability of the substance; (2) grinding the ammonium perchlorate to reduce particle size further increases thermal stability; and (3) unsaturated polymers such as the polybutadienes can be successfully used in a heat-sterilizable propellant system. Among the topics considered by the study are oxidizers, dopants, binders, and the thermal cycling of 70 lb and 600 lb propellant grains.
The ballistics behavior of nitramine propellants containing RDX in an inert binder was examined. It is shown that at 2000 to 5000 psi, depending on the particle size, a slope break or discontinuity in the log burning rate versus log pressure curve occurs. It is shown that at higher pressures the nitramine decomposition proceeds predominately by C-N bond rupture and gives CH2 and N20. At still higher pressures (after the slope break) the decomposition proceeds by N-N rupture. The decomposition of RDX was investigated from 170 to 800 C at atmospheric pressure. The major decomposition products were CH2, CO, CO2, N20, N2, and H20. The ratio of the products varied with the pyrolysis rate and temperature.
The propulsion performance and reconstructed mass properties data from Thiokol's RSRM-13 motors which were assigned to the STS-41 launch are presented. The SRM propellant, TP-H1148, is a composite type solid propellant, formulated of polybutadiene acrylic acid acryonitrile terpolymer binder, epoxy curing agent, ammonium perchlorate oxidizer, and aluminum powder fuel. A small amount of burning rate catalyst (iron oxide) was added to achieve the desired propellant burn rate. The propellant evaluation and raw material information are also presented. The presented ballistic performance was based on the Operational Flight Instrumentation. The adjustments made to the raw data on this flight include biasing the data to correct ambient pressure before liftoff. The performance from each motor as well as matched pair performance values were well within the CEI Specification requirements.
The propulsion performance and reconstructed mass properties data from Thiolol's RSRM-11 motors which were assigned to the STS-35 launch are contained. The Thiokol manufacturing designations for the motors were 360W011A/360W011B, which are referred to as RSRM-11A and RSRM-B, respectively. The launch of STS-35 occurred on 2 December 1990 at the Eastern Test Range (ETR). The data contained herein was input to the STS-35 Flight Evaluation Report. The SRM propellant, TP-H1148, is a composite type solid propellants, formulated of polybutediene acrylic acid, acryonitrile terpolymer binder, epoxy curing agent, ammonium perchlorate oxidizer, and aluminum powder fuel. A small amount of burning rate catalyst (iron oxide) was added to achieve the desired propellant burn rate. The propellant evaluation and raw material information for the RSRM-11 are included. The ballistic performance presented was based on the Operational Flight Instrumentation (OFI) 12.5 sample per second pressure data for the steady state and tail off portion of the pressure trace. Recent studies have shown that the transducers are affected by the measuring system at KSC and temperature gradients created by the igniter heaters. Therefore, an adjustment to the data from each transducer is made to make the initial reading match the atmospheric pressure at the time of launch.
The propulsion performance and reconstructed mass properties data from Thiokol's RSRM-9 motors which were assigned to the STS-36 launch are presented. The SRM propellant, TP-H1148, is a composite type solid propellant, formulated of polybutadiene acrylic acid acryonitrile terpolymer binder (PBAN), epoxy curing agent, ammonium perchlorate oxidizer and aluminum powder fuel. A small amount of burning rate catalyst (iron oxide) was added to achieve the desired propellant burn rate. The propellant evaluation and raw material information for the RSRM-9 are included. The propellant grain design consists of four segments. There is a forward segment with an eleven point star with a transition into a tapered circular perforated (CP) configuration. There are two center segments that result in a double tapered CP configuration and an aft segment with a triple taper CP configuration and a cutout for the partially submerged nozzle. The ballistic performance presented is based on the Operational Flight Instrumentation (OFI) 12.5 sample per second pressure data for the steady state and tail off portion of the pressure trace. No high sample rate pressure gauges, Development Flight Instrumentation (DFI), were used on this flight and therefore no ignition data is presented.
Contributions to the evolution of solid-propellant rocketry have come from a variety of sources. World War II research on large solids enabled one company to capitalize on work in the area of castable double-base propellants. Separate development of castable composite propellants led to production of Polaris and Minuteman powerplants. Pivotal to the development of these missiles were Edward Hall's advocacy of the Minuteman missile within the Air Force and contract funding to resolve problems. The discovery that adding large amounts of aluminum significantly increased the specific impulse of a castable composite propellant further aided large-missile technology. These separate lines of research led to the development of large solid-propellant motors and boosters. Many more discoveries went into the development of large solid-propellant motors. Ammonium perchlorate replaced potassium perchlorate as an oxidizer in the late 1940's, and binders were developed. Discoveries important in the evolution of large solid-propellant motors appear to have resulted from innovators' education and skills, an exposure to contemporary problems, an awareness of theory but a willingness not to let it dictate empirical investigations, and proper empirical techniques. Other important contributions are the adequate funding and exchange of information. However, many questions remain about these and other innovations.
Surface flame spreading characteristics of ICRPG reference composite propellant composed of ammonium perchlorate with polyurethane binder
Outgassing rate measurements on stored composite fuels of polybutadiene acrylic acid and polyurethane binders
This work was conducted to define further the effects of propellant composition variables on the acceleration-induced burning rate augmentation of solid propellants. The rate augmentation at a given acceleration was found to be a nonlinear inverse function of the reference burning rate and not controlled by binder or catalyst type at a given reference rate. A nonaluminized propellant and a low rate double-base propellant exhibited strong transient rate augmentation due to surface pitting resulting from the retention of hot particles on the propellant surface.
Structural behaviour of viscoelastic materials application to failure criteria of solid- propellant rocket motor grains
Solid propellants based on low cost hydroxyl terminated polybutadiene binders meeting burning rate, mechanical properties and processing requirements for large booster motors
Oxidizer particle size and binder type effects on nonacoustic combustion instability of solid propellants
Solid propellant extinguishment by rapid depressurization, investigating motor configuration effect, binder, burning rate catalyst, oxidizer, metal loading and exhaust pressure levels
This paper presents several improvements to the BDP model of steady-state burning of AP composite solid propellants. The Price-Boggs-Derr model of AP monopropellant burning is incorporated to represent the AP. A separate energy equation is written for the binder to permit a different surface temperature from the AP; this includes an analysis of the sharing of primary diffusion flame energy, and correction of a BDP model inconsistency in treating the binder regression rate. A method for assembling component contributions to calculate the burning rates of multimodal propellants is also presented. Results are shown in the form of representative burning rate curves, comparisons with data, and calculated internal details of interest. Ideas for future work are discussed in an Appendix.
The objective of this work was to better define the effects of propellant composition variables on the acceleration induced burning-rate augmentation of solid propellants. The rate augmentation of an aluminized propellant at a given acceleration level was found to be a nonlinear inverse function of the reference or static burning rate and not controlled by binder or catalyst type at a given reference rate. A nonaluminized and a low rate double-base propellant exhibited strong transient rate augmentation due to surface pitting resulting from the retention of hot particles on the propellant surface.
The feasibility of reducing troublesome nozzle blockage (by condensation deposits) in laboratory-scale solid rockets by addition of a silicone oil as a propellant ingredient was explored experimentally. An aluminized composite propellant and its counterpart with 1% silicone oil replacing part of the binder were fired in a 63.5 mm diameter, end-burning, all-metal burner. Pressure-time histories were recorded for all of the tests by a Taber gauge mounted at the downstream end of the chamber; temperature-time data at the nozzle throat were obtained in some of the runs by thermocouples having junctions positioned at the wall but insulated from the metal. Deposition of condensables on the nozzle walls causing a progressive increase in the chamber pressure with time was noted. The fraction of firings exhibiting practically no condensation was 59% with silicone and 32% without. On the average, temperature readings at the nozzle throat were higher with the silicone propellants. Although various phenomena may contribute to these findings, the results are not understood completely.