Rocket Motor Ignition Research Articles

Novel method for Investigation the mass of different pyrotechnics MTV Compositions for rocket motor igniter

Novel method for Investigation the mass of different pyrotechnics MTV Compositions for rocket motor igniter

Photoflash and laser ignition of Al/PVDF films and additively manufactured igniters for solid propellant

Solid propellants are employed in a range of applications from the inflation of airbags to propulsion systems for rockets. The ignition of solid propellants must be carefully controlled and modified on a per-use basis due the specific ignition requirements of each application. Using tailored photoreactive materials as a source of ignition for solid propellants, or other energetic materials, could reduce the added weight and risk of traditional initiators and result in safer, more effective solid rocket motor ignition systems. This study demonstrates the tunability of the ignition delay and propagation properties of optically-sensitive, nearly full density reactive aluminum/polyvinylidene fluoride (Al/PVDF) films and additively manufactured igniters. A single printed layer of pure nano-aluminum (nAl) at ideal stoichiometry in PVDF was found to flash ignite, but frequently yielded delayed transitions in steady propagation from the igniter to the propellant. To improve the continuity and steadiness of the transition, fuel particle size, igniter thickness, and a combination of layers of nAl and micron-sized aluminum (μAl) were investigated. In printed igniters with layers of μAl, only a single layer of nAl was needed to flash ignite the material and propagate to the layers of μAl without delay. For igniters cast onto strands of ammonium perchlorate composite propellant, continuous ignition was achieved with a single layer of nAl printed atop a triple layer of μAl for the flash-activated igniters and a single layer of nAl printed atop a single and triple layer of μAl for laser-driven igniters. The nAl/PVDF layer enabled good flash or laser ignition sensitivity, while the μAl/PVDF produced more sustained heat transfer to produce a reliable ignition process.

A Model for Igniter Mass Flow Rate History Evaluation for Solid Rocket Motors

This paper describes a zero-dimensional model for evaluating the mass flow rate history of a solid rocket motor igniter. Based on the results of an igniter-firing experiment, in which the igniter is the only source of combustion gas and no propellant is ignited, the proposed model can be used to compute the mass flow rate of the igniter. Different species and temperature-dependent properties, such as the specific heat for each species, are considered. The coupling between the flow field variables in the combustion chamber and the heat transfer at the gas-solid interface is computed in a segment way. Calculations are performed for different species and properties, and the errors are discussed. Using the computed igniter mass flow rate as a boundary condition, a two-dimensional calculation is performed for validation purposes. The results are in good agreement with experimental data. The proposed model can be used to provide reasonable boundary conditions for solid rocket motor simulations and to evaluate the performance of igniters. Although derived on the basis of a small-scale solid rocket motor, the model has the potential to be used in large-scale systems.

Dimension-adaptive algorithm-based PCE for models with many model parameters

Purpose To present the models with many model parameters by polynomial chaos expansion (PCE), and improve the accuracy, this paper aims to present dimension-adaptive algorithm-based PCE technique and verify the feasibility of the proposed method through taking solid rocket motor ignition under low temperature as an example. Design/methodology/approach The main approaches of this work are as follows: presenting a two-step dimension-adaptive algorithm; through computing the PCE coefficients using dimension-adaptive algorithm, improving the accuracy of PCE surrogate model obtained; and applying the proposed method to uncertainty quantification (UQ) of solid rocket motor ignition under low temperature to verify the feasibility of the proposed method. Findings The result indicates that by means of comparing with some conventional non-invasive method, the proposed method is able to raise the computational accuracy significantly on condition of meeting the efficiency requirement. Originality/value This paper proposes an approach in which the optimal non-uniform grid that can avoid the issue of overfitting or underfitting is obtained.

Thermokinetic Analysis and Performance Evaluation of Guanidinium Azotetrazolate Based Gas Generating Composition for Testing of Solid Rocket Motor Nozzle Closures

AbstractA gas generating composition (GGC) was designed with Guanidinium azotetrazolate (GZT) as fuel and Ammonium perchlorate (AP) as an oxidizer for burst testing of nozzle closure at rig level, simulating solid rocket motor ignition transient and propellant flame temperature. EXPLO5 software code was used for formulation studies. The selected GGC formulation was subjected to thermokinetic analysis for understanding the thermal decomposition behaviour as well as evaluating kinetic and thermodynamic parameters of decomposition.The non‐isothermal kinetic parameters of the composition over the extent of conversion were determined using two isocoversional methods viz; Flynn‐Wall‐Ozawa (FWO) and Kissinger–Akahira–Sunose (KAS). GGC in pellet form was fired using a pyrotechnic pressure generating device (PPG) in test rig level with nozzle closure. The nozzle closure was burst opened at a pressure of 0.985 MPa with a pressurization rate of 10.3 MPa s−1.

Uticaj veličine čestica pripalne smeše na gradijent pritiska pripale raketnog motora

The paper discusses the influence of ignition mixture granulation onto the regime of rocket motor ignition. Igniter tests were performed in manometric bomb and pressure distribution was measured. In each test different granulations of the same mixture in powder form were tested. After that, two different ignition mixture compositions in the form of pressed tablets were tested. Total masses of ignition mixtures were equal in all the tests. In the first series of the igniter tests, very carefully chosen different granulation of the powder was used in each test, with the particles in a narrow range. From one to another test, only the granulation of the mixture was changed. The purpose was to determine if there exists an influence of the mixture granulation onto the pressure gradient and thus also onto the regime of the rocket motor ignition. It turned out, surprisingly, that there were no visible differences in the regimes of pressure increase due to the different granulation of the ignition mixtures. In another series of tests the mixture was prepared in the form of pressed tablets, with various sizes from test to test. Two different mixtures were tested, and the mass of each single charge of the igniter was the same as in the first series of tests. It is shown that the form of the pressed tablets, as well as their composition, may considerably affect the mode of rocket motor ignition.

Three-dimensional numerical and experimental studies on transient ignition of hybrid rocket motor

This paper presents transient simulations and experimental studies of the ignition process of the hybrid rocket motors (HRMs) using 90% hydrogen peroxide (HP) as the oxidizer and polymethyl methacrylate (PMMA) and Polyethylene (PE) as fuels. A fluid-solid coupling numerically method is established based on the conserved form of the three-dimensional unsteady Navier-Stokes (N-S) equations, considering gas fluid with chemical reactions and heat transfer between the fluid and solid region. Experiments are subsequently conducted using high-speed camera to record the ignition process. The flame propagation, chamber pressurizing process and average fuel regression rate of the numerical simulation results show good agreement with the experimental ones, which demonstrates the validity of the simulations in this study. The results also indicate that the flame propagation time is mainly affected by fluid dynamics and it increases with an increasing grain port area. The chamber pressurizing process begins when the flame propagation completes in the grain port. Furthermore, the chamber pressurizing time is about 4 times longer than the time of flame propagation.

Reliability of solid rocket propellant grain under simultaneous action of multiple types of loads

A case-bonded solid propellant rocket grain is subjected to many stress-inducing loads during the service life, due to temperature, extended polymerization, transportation, vibration, acceleration, aerodynamic heating etc. and finally due to the operating pressure in the rocket motor. Composite propellant is a viscoelastic material whose mechanical properties highly depend on temperature and strain rate and sometimes may vary in the range of use of rocket motors for several orders of magnitude. Relationships between stresses and strains are much more complex than for the elastic material. Therefore, the stress and strain analysis and estimation of safety factor under the action of each individual load is quite complex and sometimes impossible. An even greater problem occurs when multiple different types of loads act simultaneously. An extreme case occurs in the moment of rocket motor ignition. Then, the very fast load acts due to the pressure, at which the propellant tensile strength is high. At the same time, the very slow thermal load acts on the grain, and in these conditions the propellant tensile strength is low. The vector addition of the stresses and strains due to different loads is not possible. It is also not possible to define the equivalent or resultant values of tensile strength and allowable strain. The principle of adding the current damage is applied here, similar to the model of cumulative damage. In addition, due to the large variations in mechanical properties of the rocket propellant, it is necessary to apply the methods of mathematical statistics for assessing the propellant grain reliability and service life.

Understanding Ignition Over Pressure During Start Up of Typical Scaled Down Solid Rocket Motors

The Ignition Over Pressure (IOP) is an unsteady pressure wave generated by the ignition of solid rocket motor during launch vehicle lift-off. This wave behaves as a blast or a shock wave followed by a low frequency excitation characterized up to 40 Hz, would cause severe damage to the launch vehicle, its structures and surroundings. However, in case of huge propulsion system, having two solid rocket motors as its boosters, due to the skew in their ignition, during lift-off, the phase difference of the wave from one side of the vehicle to the other could cause a severe moment, which is detrimental to the vehicle. The present paper deals with the occurrence of such blast waves during the testing of various scaled down solid rocket motors. Also, an attempt has been made experimentally to characterize and understand the propagation of the IOP wave causing unsteady pressure oscillations and transient pressure rise in the vicinity of a solid rocket motor. The spectral and directional characteristics of the IOP wave are also highlighted. Typical scaled-down solid rocket motors with and without nozzle shutters have been tested in horizontal firing configuration and the results are compared to study their effect. The resulting shock wave propagation has axial downstream as well as angular directivities. The pressure rise rate in the chamber is found to be directly correlated to the over pressure measured at various locations at the downstream of the nozzle.

Ignition theory investigation and experimental research on hybrid rocket motor

A transient ignition model is established based on the theoretical analysis of hybrid rocket motor ignition process. The ignition process can be divided into four stages: inert heating, ignition, flame propagation, and rapid pressure buildup. The inert heating up time takes up the main part of the ignition delay. An experiment system is designed for 90% hydrogen peroxide laboratory-scale hybrid rocket motor with catalytic ignition. Firing tests are carried out with polymethyl methacrylate (PMMA) and polyethylene (PE) fuels. The ignition criteria of 90% H2O2/PE hybrid rocket motor are established by comparing the theoretical calculations with experimental results. Comparison results show that the ignition process is governed by both the temperature criterion (TC) and the oxidizer-to-fuel ratio criterion (OFC). OFC determines ignition delay at a low oxidizer mass flux and TC determines ignition delay at a higher mass flux. In addition, the ignition delay time is more sensitive to the initial oxidizer temperature than the initial fuel temperature.

ENHANCEMENT OF COMPOSITE PROPELLANT IGNITION CHARACTERISTICS BY SURFACE ABRASION

During curing of composite solid rocket motor propellants, the propellant polymer tends to accumulate as a thin film adjacent to the mandrel used to form the internal bore of the propellant grain. This binder-rich layer coats the ammonium perchlorate crystals on the propellant surface and may inhibit flame propagation during rocket motor ignition. Poor flame propagation during ignition can affect both ignition performance and the reliability characteristics. Abrasion of the solid propellant bore surfaces to remove the binder-rich layer and expose the ammonium perchlorate oxidizer is a common procedure used to enhance rocket motor ignition. Abrasion of the propellant bore was evaluated as a means to enhance the ignition performance and reliability characteristics of Bristol 2.75-in. CRV7 rocket motors. Techniques were developed using wire brushes to scrub the surface of the propellant bore. Automated production processes were developed to provide uniform abrasion of the propellant surface, resulting in consistent enhancement of the rocket motor ignition characteristics. This paper describes the development of the CRV7 propellant bore abrasion process, and describes the results of the testing performed to quantify the effect of propellant abrasion on the rocket motor ignition characteristics.

5% Ares I scale model acoustic test: Overpressure characterization and analysis

During the ignition of a ducted solid rocket motor (SRM), rapid expansion of injected hot gases from the motor into a confined volume causes the development of a steep fronted wave. This low frequency transient wave propagates outward from the exhaust duct, impinging the vehicle and ground structures. An unsuppressed overpressure wave can potentially cause modal excitation in the structures and vehicle, subsequently leading to damage. This presentation details the ignition transient findings from the 5% Ares I scale model acoustic test (ASMAT). The primary events of the ignition transient environment induced by the SRM are the ignition overpressure (IOP), duct overpressure (DOP), and source overpressure (SOP). The resulting observations include successful knockdown of the IOP environment through use of a space shuttle derived IOP suppression system, a potential load applied to the vehicle stemming from instantaneous asymmetrical IOP and DOP wave impingement, and launch complex geometric influences on the environment. The results are scaled to a full-scale Ares I equivalent and compared with heritage data including Ares I-X and both suppressed and unsuppressed space shuttle IOP environments.

Numerical Simulation of Axisymmetric Solid Rocket Motor Ignition Transient with Radiation Effect

Numerical Simulation of Axisymmetric Solid Rocket Motor Ignition Transient with Radiation Effect

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

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. Einige Gesichtspunkte der Zeit/Temperatur-Überlagerung zur Anwendung bei der Voraussage der mechanischen Eigenschaften von Raketenfesttreibstoffen Optimale Prüfbedingungen zur Bestimmung der mechanischen Eigenschaften von Raketentreibstoffen (Temperaturen und Dehnungsgeschwindigkeitsbereiche) zur Entwicklung von Masterkurven wurden untersucht. Aus Masterkurven ist es möglich, den E-Modul, die maximale Spannung und Dehnung in weiten Bereichen der Temperatur und Dehnungsgeschwindigkeit und besonders die Bedingungen während der Zündung von Raketenmotoren vorauszusagen. Bei Verwendung von Kontrollversuchen wurde gute Übereinstimmung bei hohen Dehnungsgeschwindigkeiten mit den erhaltenen Ergebnissen aus den Masterkurven erreicht. Die Ergebnisse bei Kompositraketentreibstoffen (mit carboxyterminierten Polybutadienen, CTPB, als Binder) zeigten eine drastische Abnahme der maximalen Dehnung bei hohen Dehnungsgeschwindigkeiten und niedrigen Temperaturen. Quelques apects de la superposition temps/température appliqués à la prédiction des propriétés mécaniques de propergols solides de fusée On a étudié des conditions de test optimales en vue de déterminer les propriétés mécaniques de propergols de fusée (température et domaine de vitesse de déformation) en vue de développer des courbes principales. Les courbes principales permettent de prédire le module d'élasticité, la tension et l'élongation maximales dans de vastes domaines de température et de vitesse de déformation et, en particulier, les conditions pendant l'amorcage de moteurs de fusée. En utilisant les expériences de contrôle, on a obtenu, à des vitesses de déformation élevées, un bon accord avec les résultats provenant des courbes principales. Les résultats obtenus pour les propergols de fusée composites (avec des polybutadiènes à terminaison carboxy, CTPB, en tant que liant) ont montré une forte réduction de la déformation maximale à vitesses de déformation élevées et faibles températures.

Monte Carlo Analysis of Titan/Transfer Orbit Stage Planetary Mission Guidance System

An important part of space launch vehicle mission planning for a planetary mission is the integrated analysis of guidance and performance dispersions for both booster and upper stage vehicles. For the Mars Observer mission, an integrated trajectory analysis was used to maximize the scientific payload and minimize injection errors by optimizing the energy management of both vehicles. This was accomplished by designing the Titan III booster vehicle to inject into a transfer orbit plane and the transfer orbit stage (TOS) to correct any booster dispersions. An integrated Monte Carlo analysis of the performance and guidance dispersions of both vehicles provided sensitivities and a statistical evaluation of their guidance schemes. The polynomial guidance schemes used for the Titan III variable flight azimuth computations and the TOS solid rocket motor ignition time and burn direction derivations accounted for a wide variation of launch times, performance dispersions, and target conditions. The Titan III/TOS vehicle, which launched the Mars Observer spacecraft on Sept. 25, 1992, used these polynomial schemes. The postflight analysis indicated that a near perfect park orbit injection was achieved, followed by a hyperbolic orbit injection with less than 2cr errors.

Space Shuttle Solid Rocket Motor Testing for Return to Flight: Transient Pressure Test Article Test Program

The Transient Pressure Test Article (TPTA) test program is being conducted at a new test facility located in the east test area at the National Aeronautics and Space Administration's (NASA's) Marshall Space Flight Center (MSFC) in Huntsville, Alabama. This facility, along with the special test equipment (STE) required for facility support, was constructed specifically to test and verify the sealing capability of the redesigned solid rocket motor's (RSRM) field, igniter, and nozzle joints. The test article consists of full scale RSRM hardware loaded with inert propellant and assembled in a short stack configuration. The TPTA is pressurized by igniting a propellant cartridge capable of inducing a pressure rise rate that simulates the ignition transient that occurs during launch. Dynamic loads are applied during the pressure cycle to simulate external tank attach (ETA) strut loads present on the ETA ring. Sealing ability of the redesigned joints is evaluated under joint movement conditions produced by these combined loads since joint sealing ability depends on seal resilience velocity being greater than gap opening velocity. Also, maximum flight dynamic loads are applied to the test article which is either pressurized to 600 psia using gaseous nitrogen (GN2) or applied to the test article as the pressure decays inside the test article on the down cycle after the ignition transient cycle. This paper will present the uniqueness of this new test facility with respect to its capabilities. It will also touch on both the topic of test effectiveness versus space vehicle flight performance and new aerospace test techniques, as well as presenting a comparison between the old SRM design and the RSRM.

Ignition/duct overpressure induced by Space Shuttle solid rocket motor ignition

An analytical methodology, with a propagating solid rocket motor (SRM) exhaust front as a perturbation pressure wave generator, is developed to enhance understanding of the ignition/duct overpressure (IOP/DOP) induced during the Space Shuttle liftoff. Waveform, amplitude, and low-frequency responses of IOP/DOP are simulated. For a full-scale configuration, IOP and DOP responses are clearly isolated and distinguishable. However, in the subscale tests, an interaction region of IOP and DOP exists, which obscures interpretation of the test data. The methodology offers a first-order assessment of these phenomena with respect to the launch complex geometry and the SRM ignition transient, and provides guidance for the test data interpretation.

OPTIMIZATION OF SOLID ROCKET MOTOR IGNITER PERFORMANCE REQUIREMENTS†

A numerical technique is presented for optimizing a set of solid rocket motor (SRM) ignition control variables to achieve a specific requirement or set of requirements to be used in preliminary design of an SRM igniter. The mathematical model of the igniter transients uses a simplified ignition simulation routine to calculate igniter and SRM performance. The object function to be optimized (minimized) is typically the summation of The absolute values of the differences between the desired and computed thrust values of the SRM at specified times during ignition. Constraints are imposed in the form of limiting values of igniter characteristics, flame-spreading speeds and/or maximum rate of pressure rise in the SRM. Optimization is obtained using a direct pattern search technique to determine the required values of the controlling variables. Examples are presented which illustrate the ability of the technique to meet practical design requirements. Computational results are shown to be consistent with the static test performance of Space Shuttle SRMs. In addition.it is shown that the method is capable of evaluating certain often unknown parameters, such as flame-spreading speed, from test data.

Flame spreading phenomena in double base propellants

The rates of flame spread over the surface of a double base propellant were measured as a function of the pressure and of the oxygen concentration at 0.21 mole fraction and below, typical of conditions prevailing during rocket motor ignition. Relationships have been established between the flame spread rate and the burning rate, which show a strong dependence upon oxygen concentration and a somewhat lower dependence on the pressure. The findings can have significance in the study of aft-end ignited rocket motor transients, and also in the measurement of propellant burning rates using the strand burner technique.

A Fluidic Sounding Rocket Motor Ignition System

Fluidic sounding rocket motor ignition has been found to be feasible using a system without stored energy and with the complete absence of electrical energy and wiring. The fluidic ignitor is based on a two component aerodynamic resonance heating device called the pneumatic match. Temperatures in excess of 600 C were generated in closed resonance tubes which were excited by a free air jet from a simple convergent nozzle. Using nitrocellulose interface material, ignition of boron potassium nitrate was accomplished with air supply pressures as low as 45 psi. This paper describes an analytical and experimental program which established a fluidic rocket motor ignition system concept incorporating a pneumatic match with a simple hand pump as the only energy source.