Response Of Propellant Research Articles

Response of propellant combustion to a turbulent acoustic boundary layer

An analysis of the transitional and turbulent reactive acoustic boundary layer on a homogeneous, solid-propellant surface is conducted to investigate potential mechanisms of combustion instability. The theoretical approach utilizes a low Reynolds number turbulence closure model and a finite-difference solution procedure for an equation system of parabolic form. A new technique is developed for the condensed-phase thermal layer, in which the propellant space is mapped onto the gas space and efficiently solved using the same adaptive numerical grid. Results are obtained at an acoustic pressure node in the absence of a mean axial flow. The results indicate that acoustically induced transition can occur at relatively low acoustic pressure amplitudes, propellant response to harmonic axial velocity fluctuations is both rectified and displays a mean augmentation (D-C shift), and that nominal propellant combustion parameters lead to increasing susceptibility to acoustic transition at elevated mean pressures. As OQ Bg cp €« / H h h^ k ks L® n q R Reac Res Ret Ru r

Response of simulated propellant and explosives to projectile impact—III. Experimental and numerical results of warhead penetration and fragmentation

Summary This paper is the third of a series concerned with the effects of projectile impact on a simulated explosive or propellant, called Propergol. Experiments were conducted to study the fragmentation and perforation response of disks of this material when subjected to impact by blunt and cylindro-conical strikers. Similar tests were conducted on layered targets of Propergol and steel, and also for a simulated warhead that was struck by armor-piercing projectiles. Data were obtained by velocity measurement, high-speed photography and post-mortem target examination including collected fragments. A fragmentation oriented penetration code, AUTODYN(frag) was developed from the interfacing of a two-dimensional commercial finite difference code, AUTODYN, with a fragmentation subroutine, BFRACT, developed by other investigators. This program was utilized to study the microfracture and fragmentation processes in both monolithic and composite Propergol plates during their penetration by projectiles. In addition, numerical evaluations of the effects of simulated warhead penetration by armor-piercing bullets were conducted using the publicly available finite-element code DYNA2D. The numerical results were compared with corresponding experimental data and also with the predictions of an analytical representation of the phenomenon, described in the second paper of the series. Reasonable agreement was obtained in the domain where the hypotheses concerning the structure of the analysis and of the computations were applicable.

Response of simulated propellant and explosives to projectile impact—I. Material behavior and penetration studies

Summary As the first part of a broad investigation of the response of two propellant and explosive simulants to projectile impact, this paper first delineates their measured mechanical behavior under static and dynamic loading. Subsequently, a detailed description is given of the damage patterns generated upon penetration and perforation of homogeneous disks or composites where these disks were constrained axially by metallic plates or completely confined by the addition of a circumferential steel band. The two substances, labeled Propergol A and B, are composed of particle-embedded polymers and were found to be highly viscoelastic. Properties were determined from quasi-static, creep and relaxation and Hopkinson bar tests, as well as from acoustic wave experiments. Ballistic limits were determined both for pure Propergol and combinations of this material with adjacent metallic plates. On the basis of areal density, monolithic Propergol exhibits a substantially lower perforation resistance than steel, aluminum, or Propergol disks either prefaced or sandwiched by metallic plates. Damage patterns observed in the vicinity of the ballistic limit include radial and conical cracking, cylindro-conical (or mushroom) plugging for blunt projectiles, specimen cleavage due to slow crack propagation in constrained samples upon removal of the constraint, and debonding and rebonding in the interior around the projectile trajectory. In this velocity regime, fragmentation is minimal; at higher speeds, plug formation is inhibited and the phenomenon is fragmentation dominated. This feature and phenomenological or numerical analyses of the processes are discussed in subsequent papers.

Response of simulated propellant and explosives to projectile impact—II. Fragmentation

Summary This paper is the second of a series concerned with the penetration and perforation phenomena in two types of propellant and explosive simulant, named Propergol, due to the impact at normal incidence of both blunt and conically-tipped steel strikers. The collision results in fragmentation, plug formation and generation of a cloud of debris that includes particles of measurable dimensions traveling with significant velocities. Both the fragment size and area as well as the ejecta mass are determined experimentally as a function of Propergol specimen thickness and impact velocity or energy. The cumulative number of fragments as a function of size for the Propergol is uniformly found to be a bi-linear semi-logarithmic relationship with the bifurcation occurring at the mean crystal radius. Individual crystals and the crater generated are examined by means of a scanning electron microscope. A phenomenological model of the fragmentation process is constructed, based on an assumed spherical shape of the fragments and the bi-linear fragment distribution, using energy methods. This is combined with a perforation analysis that considers the process to be sequentially composed of initial indentation, fragmentation, and sliding and deflection of the Propergol disks. An evaluation of this model providing fragment volumes as a function of impact velocity is compared with experimental results and found to be in good agreement.

Analysis of cryogenic propellant behaviour in microgravity and low thrust environments

Understanding the behaviour and response of cryogenic propellants during spacecraft flight operations (e.g., engine restart and fluid transfer) is an extremely important aspect of vehicle design. Accurate predictions of fluid motion and slosh amplification are needed to ensure proper settling prior to engine burn and effective vehicle control throughout all phases of the mission. To augment analyses of this type, Marshall Space Flight Center (MSFC) recently acquired FLOW-3D, a CFD package developed by Flow Sciences Inc. This paper describes a recent validation in which FLOW-3D model predictions were compared with MSFC drop tower test data. Although the tests were originally conducted in the 1960s to support design and performance assessments of the Saturn S-IVB stage liquid hydrogen (LH 2) tank, the data have proven useful for verifying the accuracy of the FLOW3D model.

Assessment of the response of rocket propellants to high-velocity projectile impact using small-scale laboratory tests

Two small-scale laboratory testing methods (Hopkinson Bar and Shotgun tests) for assessing the response of rocket propellants to high-velocity projectile impact are described. The impact ignition sensitivities of several composite propellants and cast double base (CDB) propellants have been determined. Composite propellants with different mechanical properties, oxidizer loading, and oxidizer and binder types, and CDB propellants with different nitrocellulose (NC)-to-nitroglycerine (NG) ratios have been examined. The two tests give comparable results, except for very “soft” propellants. Results from the Hopkinson Bar and Shotgun tests correlate with bullet impact field trials of end-burning rocket motors and suggest that the information obtained from these small-scale tests are relevant to propellant ignition and burning and can predict, to some extent, the response of rocket propellants to projectile impact.

Harmonic analysis of piston and pyrotechnic pulsers for T-burners

MALL samples of propellants are routinely tested in a Tburner apparatus to determine their tendency to burn unstably. For spontaneously unstable propellants, small acoustic disturbances in the T-burner can couple with the propellant combustion, leading to the growth of a pressure wave from which a propellant response function can be determined. Some propellants, especially those containing metal, are not spontaneously unstable, so that pulsed disturbances must be introduced into the burner chamber during burning and at burnout to excite the pressure waves. The decay rates of these waves can then be used to determine the propellant response

The acoustic boundary layer in porous walled ducts with a reacting flow

An analysis of the acoustic boundary layer in a porous walled duct containing a primixed, laminar flame adjacent to the duct walls is presented. The reactants are injected through the sidewalls and an overall one step chemical reaction is assumed. The theory finds applications in studies of propellant response and stability analyses of solid propellant rocket motors. The formulation, valid for low injection Mach numbers, is two dimensional and includes effects of viscosity, thermal conductivity, axial and transverse velocities and unsteady combustion in the flame. It is shown that increasing the flow injection rate increases the boundary layer thickness at a given frequency and its tendency to damp acoustic motions in the duct also increases. However, the unsteady reaction rate and associated unsteady heat release rate tend to enhance acoustic motions within the duct and generally dominate the damping processes. Furthermore, the analysis shows that the unsteady expansion in the unsteady reaction region is the, mechanism responsible for the observed acoustic driving. Finally, the analysis shows that the acoustic admittance at the boundary layer edge depends upon axial location within the duct, varying most rapidly near an acoustic pressure minimum.

Flow turning and admittance correction - An experimental comparison

The proper Way of incorporating viscosity-related acoustic losses into solid propellant rocket motor linear stability analyses has been an open question for years. Mainly, two distinct theories are proposed in the literature and are referred to as the flow turning and admittance correction approaches. The two theories are briefly presented, then competitively used to derive composite solid propellant pressure-coupled response functions from experiments. The experimental setup is succinctly described and experimental results are presented. Comparisons between the theoretical approaches are made on the basis that the proper approach should yield equal values for the propellant response function, irrespective of particular grain geometry, provided that the same mean chamber pressure and frequency are achieved in the chamber. Results favor the admittance correction approach, as the flow turning approach seems to overestimate viscous losses.

Sound generation by turbulence in simulated rocket motor cavities

The present investigation is motivated by vibration problems in solid-propell ant rocket motors. A class of interior flows modeled to simulate flow conditions inside rocket motor cavities is considered. The turbulence generated pressure fluctuation consists of two components, acoustic and hydrodynamic. The Bernoulli enthalpy theory of aeroacoustics is employed to extract acoustic pressure spectra from experimentally obtained turbulence data and acoustic impedance values at flow boundaries. The effects of turbulence intensities, sidewall acoustic impedance, length-to-diameter ratio of the cavity, and different mass flux on the acoustic pressure level are investigated in experimental configurations. Typical pressure levels inside rocket motor environments are calculated utilizing the A-B representation for propellant response.

Investigation of solid propellant combustion instability by means of a T-burner

Experimental investigation of solid propellant combustion instability using variable area T-burner is presented. The test facility is described and theoretical background and formulation related to the specific geometry are given. A number of solid propellants were tested over pressure oscillation frequency range of 350–1000 Hz. Results show complex behavior of the propellant response, and can characterize the different propellants according to their combustion instability level.

The Effect of Nonlinear Heat Conduction on the Pressure-Coupled Response of Solid Propellants

A ZELDOVICH-NOVOZHILOV model for the pressurecoupled response of solid propellants was developed with heat conduction in the condensed phase governed by a relaxation time, non-Fourier relation. Results indicate that, although the relaxation time appears in a frequency-squared relaxation time product combination, the non-Fourier model has negligible effect on the response in the frequency domain of interest to combustion instability for reasonable values of relaxation time. For large values of this parameter, the resonance is shifted to lower frequencies, reduced in magnitude, and narrowed. Presented as Paper 82-1102 at the AIAA/SAE/ASME 18th Joint Propulsion Conference, Cleveland, Ohio, June 21-23, 1982; submitted June 25, 1982; synoptic received July 5, 1983. Copyright © American Institute of Aeronautics and Astronautics, Inc., 1982. All rights reserved. Full paper available from AIAA Library, 555 W. 57th Street, New York, N.Y. 10019. Price: Microfiche, $4.00; hard copy, $8.00. Remittance must accompany order. * Graduate Student, School of Aeronautics and Astronautics (currently, Engineer, Morton Thiokol, Inc., Huntsville, Ala.). Student Member AIAA. tSenior Researcher, School of Aeronautics and Astronautics. Associate Fellow AIAA.

A constitutive formulation for high-elongation propellants

A constitutive model is presented that appears to successfully represent the stress-strain behavior of a highelongation propellant over a wide range of conditions. The model uses a Lagrangian formulation and separates the propellant response into shearing and volumetric components. Particular features of the model are the use of a strain-softening function multiplying a convolution integral that is analogous to that of linear viscoelasticity. Within the integral is an additional term that modifies the history effects. This term appears to correlate behavior at changing strain rate with that at changing temperature. Comparisons with experimental data are given.

Application of the impedance tube technique in the measurement of the driving provided by solid propellants during combustion instabilities

This paper describes the adaptation of the impedance tube technique for the measurement of solid propellant response functions. The experimental setup consists of a tube with a disc-shaped solid propellant sample placed at one end and an additional propellant sample at the side wall of the tube at some distance downstream of the end propellant sample. When the pressure coupled response function is measured only the disc-like sample of the propellant is needed while determination of the velocity coupled response function requires the use of both propellant samples. During an experiment, an acoustic driver located at the opposite end of the tube is utilized to establish a standing wave of a desired frequency within the impedance tube and an array of pressure transducers mounted at preselected locations along the impedance tube walls is utilized to measure the continuously varying wave structure within the tube throughout the course of the experiment. Upon the completion of the experiment, the measured data together with solutions of the impedance tube wave equations are utilized to determine the propellant response function. The paper discusses the different data reduction procedures that have been developed for the determination of both the pressure and velocity coupled response functions. Furthermore, the paper describes the results of pressure coupled response measurements for both aluminized and nonaluminized propellants and the results of a theoretical study that show that the impedance tube method could also be used in the measurement of the velocity coupled response function.

Pressure-coupled response of composite solid propellants

Pressure-coupled response of composite solid propellants

Experimental Study on Oscillatory Combustion in Solid-Propellant Motors

The relationship between the longitudinal mode oscillatory combustion phenomena in solid-propellant motors and the motor configuration is experimentally examined using small test motors with various grain configurations. The test results show that the establishment of longitudinal oscillation closely relates to a change of the port area along the motor axis, the internal gas flow conditions, and the motor slenderness. The effects of these parameters on the motor design criterion are qualitatively discussed in relation to the propellant response function and the agglomerative burning of aluminum particles. The test results of an effective mechanical suppression device applied to an actual sounding rocket motor are also presented.

Réponse des Propergols Sol ides aux Oscillations de Pression et de Vitesse

Abstract The stability of a rocket motor depends on the response of the solid propellant to pressure and velocity fluctuations on the combustion surface, and the determination of this response is at present only possible by experimental means. The technique developed for the study of pressure coupling consists in creating, by modulation of the exhaust throat, forced pressure oscillations in the combustor of a test engine with short residence time. Progress obtained in analyzing this type of engine and in experimental procedure yielded encouraging results. The study of velocity coupling, approached with similar means, is still at the explorating stage.

Combustion Response Function Measurements by the Rotating Valve Method

A study has been conducted to evaluate the rotating valve method of measuring the combustion response of solid propellants to small amplitude pressure oscillations. The method is based on producing pressure oscillations in a small rocket motor by varying the area of a secondary exhaust nozzle in a periodic manner. This is accomplished by using a rotating valve as the secondary orifice. The valve apparatus operates concurrently with a primary nozzle which controls the steady-state pressure. The frequency of the oscillations is determined by the rotational speed of the valve. A theoretical analysis was conducted to relate the combustion response function to measurable ballistic properties of the combustion chamber. Assuming that the combustion chamber is small in comparison to the acoustic wavelength, the combustion response function can be calculated from the amplitude of the oscillating pressure and the phase angle between the oscillating pressure and oscillating nozzle area. Cold flow tests were conducted using nitrogen and helium to test the validity of the analysis. Excellent agreement was found between the measured and predicted amplitudes and phase angles. Combustion tests then were conducted using two aluminized propellant formulations and three nonaluminized formulations. There was excellent agreement between the T-burner and rotating valve tests conducted on the same batch of propellant. For two nonaluminized propellants, the comparisons were based on different batches of propellant. Differences in combustion response and variations in burning rate and characteristic exhaust velocity were observed for these two formulations. It was concluded that the rotating valve method is a valid substitute for the T-burner. Substantial reductions in the cost of characterizing propellants also were obtained using this new approach.

Solid propellant acoustic admittance corrections

The acoustic behavior of solid rocket propellants is conveniently expressed in terms ofthe specific acoustic admittance of the burning surface. Numerous measurements of the admittance have been made in T-burners, and it is generally assumed that the data can be used directly in estimates of rocket motor acoustic stability. However, as this paper shows, the propellant response to incident acoustic waves is influenced by the orientation of the propellant surface to the waves. If there is a tangential component of acoustic velocity, then viscous effects give rise to the generation of a traveling shear wave which modifies the apparent propellant response. This wave is out of phase with the velocity fluctuations and obviously represent an energy loss to the system. It is demonstrated that the shear wave is strongly influenced by the mean flow due to propellant combustion; for larger values of burning surface Mach number, the speed of propagation approaches that of the combustion gases. Thus admittance data from T-burners with the propellant samples arranged with normal wave incidence must be corrected for the shear effects if they are to be used in motor calculations wherein acoustic modes with grazing incidence are of interest. For the same reason it is essential to interpret carefully T-burner data secured in devices employing extended or cup-shaped grains, since part of the propellant produces shear waves and thus exhibits a different effective admittance. The corrections are shown to be of the order of the mean flow Mach number; that is, they are of the same order as the real part of the admittance and must not be ignored in motor calculations. It is demonstrated that inclusion of the corrections brings about a marked improvement in agreement between the classical linearized theory of acoustic combustion instability and experimental observations of the phenomenon.

ONERA and SNPE Recent Work on Solid Propellant Combustion Instability

Results are presented on the response of a solid propellant to pressure and velocity fluctuations. Two different experimental techniques have been developed. The first makes use of devices known to be unstable (L* or T rocket motors). Comparison of cold and hot tests shows the poor credibility of the results obtained with T rocket motors; furthermore the temperature heterogeneity is scale dependent. In the second technique, research rockets with a modulated throat section are used. These methods are discussed in some detail from the experimental and theoretical standpoints. Here again cold gas comparative tests emphasize the effect of combustion on the observed gain. A linearized theory predicts the observed trends.