Showing papers on "Solid-fuel rocket published in 1999"

Thermal properties measurements of solid rocket propellant oxidizers and binder materials as a function of temperature

Abstract: Using a fairly simple technique and small samples it was possible to obtain thermal diffusivity, specific heat capacity, and thermal conductivity, all as a function of sample temperature, for a variety of ingredients used in solid rocket propellants. The oxidizers AP, ADN, CL20, HMX, RDX, HNF, TNAZ were studied as well as the nonenergetic polymers TeflonTM, HTPB, and polyurethane, energetic binders containing GAP and BAMO and/or NMMO, and actual solid propellants XM39, N5, N12, and SB129.

Some Aspects of Time‐Temperature Superposition Principle Applied for Predicting Mechanical Properties of Solid Rocket Propellants

Abstract: Optimal test conditions for determining the mechanical properties of rocket propellants (temperatures and strain rates ranges) for delivering master curves were investigated. From master curves it is possible to predict the modulus, maximum stress and maximum strain in vide intervals of temperatures and strain rates, and especially the existing conditions during the ignition of rocket motor. Using the control experiments, at high strain rates, the good agreement between the results obtained from master curves was shown. The obtained results for composite rocket propellants (with carboxy-terminated polybutadiene, CTPB, as a binder), point out the drastic decreasing of maximum strain at high strain rates and low temperatures.

In situ measurement of the aerosol size distribution in stratospheric solid rocket motor exhaust plumes

Abstract: The concentration and size distribution of aerosol in the stratospheric exhaust plumes of two Space Shuttle rockets and one Titan IV rocket were measured using a two component aerosol sampling system carried aboard a WB-57F aircraft. Aerosol size distribution in the 0.01 µm to 4 µm diameter size range was measured using a two component sampling system. The measured distributions display a trimodal form with modes near 0.005 µm, 0.09 µm, and 2.03 µm and are used to infer the relative mass fractionation among the three modes. While the smallest mode has been estimated to contain as much as 10% of the total mass of SRM exhaust alumina, we find show that the smallest mode contains less than 0.05% of the alumina mass. This fraction is so small so as to significantly reduce the likelihood that heterogeneous reactions on the SRM alumina surfaces could produce a significant global impact on stratospheric chemistry.

Computation of Vortex Shedding in Solid Rocket Motors Using Time-Dependent Turbulence Model

Abstract: This paper is devoted to the computation of turbulent effects on the aeroacoustic interaction in solid rocket motors. The purpose is to establish a simulation method that estimates the vortex-shedding and acoustic-wave interaction. To perform these unsteady computations, turbulent motion is taken into account by a e rst-order model based on a nonlinear relationship between time-dependent Reynolds stresses and velocity gradients. Model coefe cients are explicit functions of both strain and rotation. This model is applied to the computation of two simplie ed cone gurations of a rocket booster. To better evaluate the turbulent effects, computations are presented with and without a turbulence model. In both cone gurations, the natural unsteadiness of the e ow is captured and the comparisons with numerical and experimental data are presented. The interaction between vortex shedding and acoustics is described and the turbulence effects are characterized.

Aluminum combustion modeling in solid propellant environments

Abstract: An improved numerical model of aluminum particle combustion has been developed based on previous modeling work done at Brigham Young University. Thermochemical calculations have been performed to determine the range of temperatures, pressures, and species distributions resulting from the combustion of various formulations of rocket propellants. The previouslydeveloped model has been modified to take into account such species distributions at high temperatures and pressures. Kinetic rate equations are introduced for the homogeneous reaction of COZ and Hz0 with ahuninum The model has produced results which are consistent with experimental data from many authors. In addition, the model has been used to predict alumirmm combustion rates for the range of conditions predicted by thermochemical calculations to be present inside the rocket motor, for which there is little or no experimental data. The effects of less reactive species such as HCl and HZ have been included for pressures up to 70 atm and temperatures up to 4000 K. Introduction Aluminum powder is presently used, and has been used for many years, as an ingredient to increase the specific impulse of solid rocket propellants because of the large amount of heat generated during the aluminum oxidation reaction. This occurs when aluminum reacts with the available Hz0 and CO, (the major oxidizing species) produced from a burning solid propellant In solid propellants, aluminum is used in quantities of lo-20% by mass, and the particles are typically 20-30 microns in diameter. During heat-up, these particles may melt and coalesce into larger agglomerates, ranging from 100-200 microns in size. It is very useful to be able to predict the time required for the aluminum particles to burn once they are ignited. Empirical correlations may be used, but data collected from rocket motors themselves are virtuahy non-existent because of the harshness of the motor environment Most empirical correlations have therefore been derived from carefully controlled lab experiments. However, these experiments differ greatly from the conditions in a rocket motor, and while such experiments are informative, it can be difficult to extrapolate their results to the conditions of a rocket motor. Computer modeling of ahnninq combustion thus becomes an attractive alternative. Background When aluminum ignites, the heat.of reaction is so great that the aluminum boils and thus remains at about 2800K (at 1 atm). The outward flux of gaseous ahuninum causes a flame zone to form at about 2-4 times’ the diameter of the particle, in a manner similar to that of a burning hydrocarbon droplet. In this thune zone a homogeneous reaction takes place between the ahnninq and, available oxidizer(s). Figure 1 shows a representation of the vaporphase aluminum combustion process. In the flame zone, the oxidized products consist of gas-phase AlxO, species such as AlO. Figure 1. Representation of the vapor-phase aluminum combustion process.’ There is no gas phase Al203 since Al203 dissociates at its boiling point However, there is usually not enough heat for all the oxide species to remain in the gas phase, so some A&O, species condense and associate to form

Variable burn-rate propellant

Abstract: Disclosed are propellants such as may be used in solid rocket motors. In one preferred embodiment, the propellant comprises one high energy propellant composition comprising a homogeneous mixture of fuel and oxidizer having a predetermined fuel/oxidizer ratio, wherein individual fuel particles are generally uniformly distributed throughout a matrix of oxidizer, and a low energy propellant comprising a fuel and oxidizer. The amounts of the two propellants are present in amounts which achieve a preselected burn rate.

Development of Thermal Barriers for Solid Rocket Motor Nozzle Joints

Abstract: The Space Shuttle solid rocket motor case assembly joints are sealed using conventional 0-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) 0-rings and polysulfide joint-fill compound. In the current design, 1 out of 7 motors experience hot gas to the wiper 0-ring. Though the condition does not threaten motor safety, evidence of hot gas to the wiper 0-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 performed 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 minutes, three times longer than 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 of up to a 2800 and 2560 F. respectively. Gas temperature 1/4" downstream of the thermal barrier were within the downstream Viton 0-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 second 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 0-rings.

Cryogenic solid hybrid rocket engine and method of propelling a rocket

Abstract: A cryogenic solid hybrid engine with a solid propellant chamber, a first propellant within such chamber in which the first propellant is in solid form in the chamber and is in fluid form at room temperature, a coolant fluid chamber and a coolant fluid in the coolant fluid chamber being maintained at a temperature blow the freezing point of the first propellant. The invention also relates to a method for propelling a rocket and a method for forming a solid propellant grain for use in a cryogenic solid hybrid rocket engine.

Modeling the structure of composite solid rocket fuel

Abstract: Results are reported on the creation of a model for the structure of a composite solid rocket fuel in terms of such concepts as “pockets” and “interpocket bridges.” Appropriate model calculations and experimental data are presented, and the possibility of using the model for predicting the characteristics of the combustion process is pointed out.

Solid rocket propellant

Abstract: A solid propellant composition includes an oxidizer, a fuel and a binder, the oxidizer containing a significant amount of bismuth oxide (Bi 2 O 3 )

Temperature and Pressure Sensitivities of Burning Wave Parameters of Nitramine-Containing Propellants and HMX

Abstract: Burning rate and burning surface temperatures have been obtained experimentally for modern double-base propellants containing nitramine additives and for HMX. Temperature sensitivities of burning wave parameters were estimated from the experimental data measured at pressures of 20 and 100 atm and at sample temperatures of i 80, +20, +100 C. Evaluation of standard deviations of the estimations shows that smoothing procedures for dependencies of burning wave parameters on pressure and temperature are necessary to obtain correct values of thesensitivities. Calculations ofthecriteria forstable combustion show that Novozhilov’ s [Novozhilov, B. V., “ Nonstationary Combustion of Solid Rocket Fuels,” Nauka, Moscow, 1973 (Translation AFSC FTD-MD-24-317-74) (in Russian) and Novozhilov, B. V., “ Theory of Nonsteady Burning and Combustion Stability of Solid Propellants by Zel’ dovich‐ Novozhilov Method,” Nonsteady Burning and Combustion Stability of Solid Propellants , edited by L. DeLuca, E.W. Price,and M. Summere eld, Vol. 143,Progressin Astronauticsand Aeronautics,AIAA,Washington, DC, 1992, Chap. 15, pp. 601‐ 641] criterion correctly predicts regimes of stable combustion of the propellants and HMX. An analysis of burned surface irregularities of the propellants cone rms one-dimensional character of the combustion under the investigated conditions.

Simulation of water injection into a rocket motor plume

Abstract: Within the framework of the elaboration of a plan for the preservation in operational condition of the ground trial facilities in an expertise and tests centre, a computational study on the interaction of a solid rocket motor plume with the deflector wall of a test bench, and finally with an injection of cooling water, was initiated. It was first and foremost focused on the preservation in operational condition of a new test bench specially designed for the ground trials of the solid-propellant propulsive stages of the new generation ballistic missile M51. For the study continuation we put that the motor’s propulsive gas output is X Kg/s for a 501 floor and Y for a 401 floor. The test facility is equipped with a nominal system of water injection which is such that the radial water output is equal to 40% X of the motor gas output and the parietal output is 110% X of the latter. A first study [II was achieved with the help of the computer code FLUENT4 [Zl in the middle of 1998 when, owing to the restrictive features of this code, the water injection in liquid phase had to be simulated by an injection of equivalent vapour. Then, a Copyright

Apparatus for after-burning fuel rich rocket exhaust products

Abstract: An apparatus for the disposal of solid rocket motors which produce exhaustases containing flammable and/or explosive products The solid rocket motor to be burned is first detachably connected to and inserted in one end of an elongated mixing chamber which has, at an opposite end, a multi-step expansion nozzle The rocket motor is then ignited and the exhaust gas is passed into the mixing chamber, which is sufficiently large to reduce the velocity of exhaust gases from supersonic to subsonic Air is injected into the mixing chamber to react with flammable and/or explosive products in the exhaust gases and the resultant mixture is then discharged from the mixing chamber through a multi-step expansion nozzle to expand the gaseous mixture and, at the same time, produce substantial turbulence within the gas mixture which is about six times higher than the turbulence levels produced by a constant area circular nozzle The expansion nozzle has an inner wall with a stair-step shaped surface, having an aspect ratio of not less than 2 and not greater than 10, to create the increased turbulence levels within the multistep nozzle

Modular cryogenic solid fuel rocket drive sets

Abstract: Modular, cryogenic, solid rocket propellants of different fuel components, such as fuels, oxidizers, energy-increasing admixtures, binders, additives, etc., for all applications of solid rockets produce a uniform, stable and complete combustion, This objective is accomplished since one of the fuel elements, due to the special selection of its composition, is provided as a permanent igniter generator of the modular propellant.

Active control of vortex-driven oscillations in a solid rocket motor using a cold-flow simulation

Abstract: The Aerospace Corporation has been conducting an experimental program on active control of vortexdriven instabilities in a solid rocket motor (SRM). This type of low-frequency combustion instability is due to the coupling between an internal vortex shedding frequency and a ‘natural acoustic frequency of the motor. Experiments using a new cold-flow simulation are underway. The 3-dimensional, Plexiglas model uses air as the primary fluid and has both a choked inlet and a choked exit flow to simulate the closed-end acoustics of a SRM. Two orhfice plates are responsible for the vortex shedding and impingement. The cold-flow facility produces fixed-frequency, vortex-driven pressure oscillations near the 3’d axial (closed-closed cavity) acoustic mode near 600 Hz and the I” axial mode at about 200 Hz. Frequency tracks similar to ones seen in actual hot-fire motors were demonstrated using a variable-area nozzle. The time-dependent exit area results in a changing internal Mach number and, hence. a timedependent vortex shedding frequency. This frequency changes continuously from 635 to 550 Hz and then jumps to another track and changes continuously from 270 to 250 Hz over a 7-second time period. A simplifted model of the chamber pressure response to the secondary-injection active control technique is also presented.

Launch Vehicle Dynamic and Control Effects from Solid Rocket Motor Slag Ejection

Abstract: The effects of the ejection of slag from a Titan IV solid rocket motor upgrade (SRMU) are analyzed for the maximum dynamic pressure time. It is assumed that the 2500 Ibs of slag on board at that time is ejected, and the corresponding thrust perturbation is determined using existing theory for the discharge of an inert body. The slag ejection event causes sharp thrust transients during blockage of the nozzle throat and blow-down of the motor internal pressure. Control simulations indicate a worst-case SRMU gimbal angle change of 0.51 deg during the event, and small-amplitude limit cycles after the excitation. An aeroelastic analysis yields yaw shear loads and bending moments as high as 13600 Ibf and 5.1x 106 Ibf-in, respectively, at the aft end of the vehicle. It is shown that the degree of the nozzle throat blockage primarily defines the lateral load magnitudes. The Titan IV has adequate structural and control margins, and slag ejection does not pose a concern for its missions. Nomenclature A* =

Control of Vortex-Induced Oscillations in a Model Solid Rocket Motor: Flow Characterization and Secondary-Injection Experiments

Abstract: : Vortex-driven instability in a solid rocket motor and its active control via secondary injection were studied in a cold-flow experimental model. All tests were conducted in an 86-cm-long, 5- cm-square airflow chamber. Twin orifice plates at the center of the chamber produced a 21 5-Hz vortex-driven pressure instability with 2% peak-to-peak oscillations near the first longitudinal chamber mode. Hot-wire anemometer data at the vortex shedding frequency indicate that the organized structures occur between the two orifice plates located at the center of the chamber and not between the orifice pair and the exit nozzle. Exploratory secondary-injection active control experiments were performed by pulsing a fast-acting solenoid valve at the upstream end of the chamber. The chamber pressure response (without orifice plates) to the pulsating secondary injection was characterized for various flow rates at frequencies of 100 and 150 Hz. Pressure perturbations of sufficient magnitude (1-5% peak-to-peak)to counter the vortex-driven instabilities were demonstrated. An active control scheme utilizing the present secondary-injection hardware will be the subject of future studies.

Investigations on Liquid Fly-Back Boosters Based on Existing Rocket Engines

Abstract: This paper describes a parametrical investigation of different types of liquid fly-back boosters (LFBB)performed at DLR. The preliminary design and performance estimation is exclusively restricted to the incorporation of rocket motors, already under development or in operation. Staged combustion as well as gas-generator cycle is looked upon. The focus is on the LOX/RP1 RD-170/172 of NPO Energomash/Pratt&Whittney, the NK-33 (AJ-26) of Kuznetzov/Aerojet, the cryogenic RD-0120 of KB Khimavtomatiki, and the LOX/LH2 Vulcain 2 of SNECMA/SEP. The attached reference expendable space transportation system is the Ariane 5 ECB with cryogenic upper stage, but skipped SRB. The basic requirement is to reach almost the same commercial GTO payload as for the launcher with solid rocket boosters. The investigation includes trajectory simulations and optimizations for ascent. Return to the launch site by the LFBB is regarded, concerning the propellant requirements and the loads on the vehicle. The booster geometry is generated by CAD for a preliminary dimensioning and mass estimation. The paper includes a comparison of size and mass, as well as flight data of the different liquid fly-back booster configurations. The relevant rocket engine figures of performance, mass, reusability, and throttling capability are presented.

Analysis of Liquid Fly-Back Booster Performance

Abstract: This paper describes in its first part a parametrical investigation of different types of liquid fly-back boosters (LFBB). The preliminary design and performance estimation is exclusively restricted to the incorporation of rocket motors already under development or in operation. Staged combustion as well as gas-generator cycle is looked upon. The focus is on the LOX/RP1 RD-170/172, the NK-33 (AJ-26), the cryogenic RD-0120, and the LOX/LH2 Vulcain 2. The attached reference expendable space transportation system is the Ariane 5 ECB with cryogenic upper stage, but skipped SRB. The basic requirement is to reach almost the same commercial GTO payload as for the launcher with solid rocket boosters. The investigation includes trajectory simulations and optimizations for ascent. Return to the launch site by the LFBB is regarded, concerning the propellant requirements and the loads on the vehicle. Different booster geometries are generated by CAD for a preliminary dimensioning and mass estimation. Critical flight stability aspects are preliminary assessed. In a second step an evolutionary parallel staged launcher of the FESTIP system study is regarded. Originally based on a fly-back booster with expendable Ariane 5 core stage, the vehicle is designed to eventually evolve to a fully reusable sytem. Preliminary calculations and lay-out was done within the system concept team of FESTIP. A succeeding study by DLR of this latter configuration is presented. Since it is based on propellant crossfeed between booster and orbiter, the throttling requirements at booster burn-out are also of interest, due to their influence on performance. The paper includes a comparison of size and mass, as well as flight data of the different liquid fly-back booster configurations. The relevant rocket engine figures of performance, mass, reusability, and throttling capability are presented. A sensitivity analysis concerning the growth factor is given for both fly-back booster applications.

Solid Rocket Motor Development for Land-Based Intercontinental Ballistic Missiles

Abstract: Solid rocket motor development in the U.S. Air Force land-based intercontinental ballistic missile programs is briee y described. The new, radical approach to program management adopted by Bernard Schriever and the U.S. AirForceWesternDevelopmentDivisionin1954ispresentedasoneofthemajorreasonsfortheoutstandingsuccess of Minuteman I, II, and III; Peacekeeper; and Small Intercontinental Ballistic Missile Program. A broad technical discussion is included of solid rocket motor technical challenges and growth in performance and capability for Minuteman through Peacekeeper and the small intercontinental ballistic missile. Particularly signie cant are the service life characteristics of the three-stage solid rocket boosters. Data and future predictions are also provided. Finally, current and future challenges in the development of solid rocket motors to sustain the U.S. Air Force deployed intercontinental ballistic missile weapon systems are briee y discussed.

Combined Structural Oscillation Effects o:n Solid Rocket Internal Ballistics

Abstract: The combined effects of radial and axial vibration of the surrounding structure on the internal ballistics of a solid rocket motor are investigated via numerical simulation. A finite-difference model is applied for the radial deformation dynamics of the propellant/casing assembly along the length of the motor, while the nonsteady internal core flow is modelled using a primarily second-order, finite-volume random-choice technique. Predicted nonsteady combustion and flow behavior resulting from an initial pressure pulse within the flow, allied to the free axial and radial oscillation behavior of the surrounding structure, is consistent with experimentally observed trends associated with axial and transverse combustion instability. Instability-related phenomena such as the dc pressure rise and axial pressure wave strength development are demonstrated to be dependent in part on the motor structural characteristics.

Sensitivity of Solid Rocket Propellants for Card Gap Test

Abstract: Card gap test, which is standardized in Japan Explosives Society, was modified in order to apply it to solid rocket propellants and carried out to evaluate sensitivities against shock stimuli. Solid propellants tested here were mainly azide polymer composite propellants, which contained ammonium nitrate (AN) as a main oxidizer. Double base propellant, composed nitroglycerin and nitrocellulose (NC), and ammonium perchlorate (AP)-based composite propellants were also evaluated in order to compare with the azide polymer propellants. It is found that the sensitivity was dominated by the oxidizer characteristics. AP-and AN-based propellant had less sensitivity and HMX-based propellant showed higher sensitivity, and the adding of NC and TMETN were contributed to worse sensitive for the card gap test. Good relationship was obtained between the card gap sensitivity and the oxygen balance of propellants tested here.

Effect of motor length and propellant formulation on nozzleless solid rocket performance

Abstract: Experimental results are presented for four nozzleless motors of different length-diameter (L/D) ratios using two different composite propellants. The experimental observations discussed are: the premature unchoking in motors of insufficient L/D ratios and the tendency for the propellant to extinguish under highly negative pressure gradient environment, both peculiar to nozzleless operation. A simple one- dimensional numerical scheme is presented to predict the performance of a nozzleless solid motor. Erosive burning, elastic grain deformation and L/D ratio-dependent combustion efficiency are considered in the scheme. A relatively simple procedure is followed to account for the coupling effect between port gas dynamics and elastic grain deformation. The experimental results are compared with those predicted by the numerical scheme. The predictions are in reasonable agreement with the experimental values. For a composite propellant the burning rate index n can significantly vary with respect to pre...

Modular solid-fuel rocket propelling charges with a casing, a fill-drain system, cooling and suspension

Abstract: The invention relates to modular solid-fuel rocket propelling charges comprising different propellant components and other components, for example, fuels, oxidators, energy increasing admixtures, binders, additives, coatings, inhibitors, etc. which can be completely or partially fragmented, i.e. they are not provided in the form of the conventional quasi homogeneous mixture, rather they are available in the form of one or more macroscopic combustible elements having any appropriate configuration. All or individual components can also be comprised of substances which must first be sufficiently brought to a solid state by cooling. The aim of the invention is to provide modular solid-fuel rocket propelling charges of the aforementioned type such that the development of dangerous situations resulting from a cooling failure during simultaneous power increases of the carrier rockets is drastically minimized. To this end, the modules are provided as an assembly or the individual combustible elements thereof are provided with a special casing (1). Said special casing permits the storage of such propelling charges and/or the operation thereof according to prescription in such a way that problems do not occur due to the dependence on temperature of the mechanical characteristics or of the state of the aggregate.