Showing papers on "Rocket published in 2004"

Green Rocket Propulsion by Reaction of Al and Mg Powders and Water

Abstract: The efficacy of using aluminum-water and magnesium-water as propellants for underwater thruster applications has been investigated by the authors. The theoretical specific impulse for both reactant systems is high, and the products of reaction (alumina, magnesia, and hydrogen) are environmentally benign. The attractiveness of these systems as “green” propellants has been commented on previously, however, no practical experimentation with these systems has been made. The present work describes the testing of a linear combustor with magnesium-water and aluminum-water under conditions of pressure and oxidizer-fuel ratios and with a metal powder feed system that could be employed in actual rocket engines. Measurements of off-design specific impulse are compared with theoretical predictions that take into account two-phase losses. Measurements of heat fluxes available to vaporize regeneratively the liquid water oxidizer are presented as well. Perhaps of most importance, observations of the degree of product oxide accumulation in the combustor are presented. These measurements and observations are used to determine the effectiveness of these two metal fuel systems as practical green propellants.

Side-Load Phenomena in Highly Overexpanded Rocket Nozzles

Abstract: The operation of rocket engines in the overexpanded mode, that is, with the ambient pressure considerably higher than the nozzle exit wall pressure, can result in dangerous lateral loads acting on the nozzle. These loads occur as the boundary layer separates from the nozzle wall and the pressure distribution deviates from its usual axisymmetric shape. Different aerodynamic or even coupled aerodynamic/structural mechanic reasons can cause an asymmetric pressure distribution. A number of subscale tests have been performed, and three potential origins of side loads were observed and investigated, namely, the pressure fluctuations in the separation and recirculation zone due to the unsteadiness of the separation location, the transition of separation pattern between free-shock separation and restricted-shock separation, and aeroelastic coupling, which indeed cannot cause but do amply existing side loads to significant levels. All three mechanisms are described in detail, and methods are presented to calculate their magnitude and pressure ratio at which they occur.

The MaCWAVE/MIDAS rocket and ground-based measurements of polar summer dynamics: Overview and mean state structure

Abstract: [1] The MaCWAVE/MIDAS collaborative rocket and ground-based measurement programs were performed at the Andoya Rocket Range and the nearby ALOMAR observatory in northern Norway during July 2002 The summer component of the MaCWAVE (Mountain and Convective Waves Ascending Vertically) program was focused on gravity wave propagation, instability, and wave-wave and wave-mean flow interaction dynamics contributing to summer mesopause structure and variability The MIDAS (Middle Atmosphere Dynamics and Structure) program concentrated on small-scale dynamical and microphysical processes near the summer mesopause Our merged program yielded a comprehensive data set comprising two ∼12-hour rocket salvoes, including 25 MET rockets and 5 sounding rockets, ground-based lidar, radar, and balloon data, and coordinated overpasses of the TIMED satellite This paper describes the measurement program and rationale, the mean state observed during the rocket salvoes, and evidence that the mean state structure during 2002 differed in important respects from previous years

The MaCWAVE/MIDAS rocket and ground-based measurements of polar summer dynamics: Overview and mean state structure : The MaCWAVE-MIDAS Program to Study the Polar Summer Mesosphere

Abstract: The MaCWAVE/MIDAS collaborative rocket and ground-based measurement programs were performed at the Andoya Rocket Range and the nearby ALOMAR observatory in northern Norway during July 2002. The summer component of the MaCWAVE (Mountain and Convective Waves Ascending Vertically) program was focused on gravity wave propagation, instability, and wave-wave and wave-mean flow interaction dynamics contributing to summer mesopause structure and variability. The MIDAS (Middle Atmosphere Dynamics and Structure) program concentrated on small-scale dynamical and microphysical processes near the summer mesopause. Our merged program yielded a comprehensive data set comprising two ∼12-hour rocket salvoes, including 25 MET rockets and 5 sounding rockets, ground-based lidar, radar, and balloon data, and coordinated overpasses of the TIMED satellite. This paper describes the measurement program and rationale, the mean state observed during the rocket salvoes, and evidence that the mean state structure during 2002 differed in important respects from previous years.

Compressible Wall-Injection Flows in Laminar, Transitional, and Turbulent Regimes: Numerical Prediction

Abstract: Numerical simulations of compressible rocket flows are conducted in laminar, transitional, and turbulent regimes. The laminar simulation is carried out on a planar rocket flow without nozzle using the unsteady two-dimensional Navier-Stokes system. The transitional and turbulent flows are performed in three-dimensional on an extended rocket geometry with a divergent outlet using compressible large eddy simulation (LES) models. In both cases, the compressibility effect plays an important role. In the laminar case, pressure oscillation is forced at the outflow boundary. The time-averaged part of the solution is compared with the inviscid theory of compressible rocket flow of Balakrishnan et al. (Balakrishnan, G., Linan, A., and Williams F. A., "Compressibility Effects in Thin Channels with Injection," AIAA Journal, Vol.29, No. 12, 1991, pp. 2149-2154) and the oscillatory part with the acoustic layer model of Majdalani and Van Moorhem (Majdalani, J., and Van Moorhem, W. K., "Improved Time-Dependent Flowfield Solution for Solid Rocket Motors," AIAA Journal, Vol. 36, No. 2, 1998, pp. 241-248). The mean flow from the present numerical result is in better agreement with the compressible theory than the conventional Taylor's profiles (Taylor, G. I., "Fluid Flow in Regions Bounded by Porous Surfaces," Proceedings of the Royal Society of London, Series A: Mathematical and Physical Sciences, Vol. 234, 1956, pp. 456-475), as expected. The oscillatory part of the flow agrees well in the first quarter of the axial extent, near the head end. Farther downstream, the discrepancies develop rapidly between the numerical result and the acoustic-layer model. Possible causes of the difference are the effect of compressibility, which alters the local speed of sound, hence, acoustic properties, and the interference of hydrodynamic instabilities. In the transition and turbulent regimes, the dynamic LES model is applied on different resolutions. The measurements data of Traineau et al. (Traineau, J. C., Hervat, P., and Kuentzmann, P., "Cold Flow Simulation of a Two Dimensional Nozzleless Solid Rocket Motor," AIAA Paper 86-1447, June 1986) are employed for comparison purposes. The refinement study by comparison with the measurement data suggests the importance of resolving the laminar and transition region for a reliable application of LES in transitional flows. With the consideration of this aspect, LES with efficient grid size can produce resonable accuracy. Forcing hydrodynamic instabilities and a more realistic injection fluctuations model are recommended.

System for interception and defeat of rocket propelled grenades and method of use

Abstract: A method for intercepting and a defeating rocket propelled grenade (RPG) which includes the steps of detecting a thermal signature from a launch of the RPG; and cueing a narrow beam radar which locates the RPG and develops a ballistic solution and target intercept point for intercepting the PPG with an intercept vehicle.

A Model for Design and Analysis of Regeneratively Cooled Rocket Engines

Abstract: A new model for the design and analysis of a regeneratively cooled rocket engine is developed. In this model two proven rocket thermal analysis codes, TDK and RTE, were conjugated. The integration of these codes was accomplished via an interface file. The accuracy of this combined TDK-RTE model was examined by comparing its results to those of other methods for the SSME and experimental data for a liquid oxygen cooled RP1/LOX engine. Several of the additions and modifications incorporated into this model make it an excellent tool for designing the cooling circuits of regeneratively cooled engines.

Time-Resolved Fuel-Grain Port Diameter Measurement in Hybrid Rockets

Abstract: A novel technique is presented for determining the instantaneous spatially averaged port diameter of solid-fuel grains in hybrid rocket motors. This technique requires measurement of the frequency of the Helmholtz oscillation of the motor and is based on the principle that this frequency is inversely proportional to the square root of the chamber volume. This technique was applied to a hybrid rocket motor burning paraffin wax with gaseous oxygen. The calculated variation of port diameter agreed well with the correlation for average regression rate, determined from mass loss during operation. A major advantage is that the only instrumentation required for implementing this technique is a high-speed pressure transducer or a photomultiplier tube.

Fuel for jet, gas turbine, rocket, and diesel engines

Abstract: A fuel or fuel blendstock for jet, gas turbine, rocket, and diesel engines, particularly jet, rocket, and diesel engines that utilizes components of conventional petroleum not currently utilized for jet, gas turbine, rocket, and diesel fuels, such as benzene, linear, and lightly branched alkanes, that may be alkylated with aromatic moieties to make monoaromatics for use in jet and diesel fuels. Additionally, a fuel having such monoaromatics having multiple desired properties such as higher flash point, low pour point, increased density, better lubricity, aerobic degradability, reduction in toxicity, and additionally can deliver benefits in blendstocks.

Scramjet Experiments using Radical Farming

Abstract: Scramjet engines are the focus of considerable interest for propulsion in the hypersonic flow regime. One of the serious technical challenges for developing scramjets is reducing the skin friction drag on the engine. The combustion chamber, in particular, is a major contributor to the skin friction drag because of the high density of the flow through that region. This investigation focuses on reducing the combustion chamber skin friction drag by minimising the surface area and size of the combustion chamber and by employing a novel approach to accomplishing combustion. The first design criterion is addressed by using a single internal-combustor scramjet configuration, as opposed to multiple external combustors, and by injecting the fuel on the intake to reduce the mixing length required in the combustor. The second design criterion refers to the use of a new technique called radical farming. This uses the highly two-dimensional nature of the flow through the engine, which is created by deliberately ingesting the leading edge shocks, to achieve combustion at lower mean static pressures and temperatures than generally expected. A simplified approximate theoretical analysis of the radical farming concept is presented. Experiments were conducted in the T4 free-piston shock tunnel on a scramjet model with a single rectangular constant cross-sectional area combustion chamber. Pressure measurements were taken along the centreline of the intake, combustion chamber and thrust surface and across the model width at three locations. Gaseous hydrogen fuel was injected halfway along the intake at a range of equivalence ratios between zero and one. The combustion chamber height was varied from 20mm to 32mm, which varied the contraction ratio of the engine from 4.1 to 2.9. The experiments were conducted at a stagnation enthalpy of either 3MJ/kg or 4MJ/kg. The nominal 3MJ/kg condition corresponds to Mach 7.9 flight at an altitude of 24km. The majority of the 4MJ/kg experiments were conducted at a nominal condition corresponding to Mach 9.1 flight at an altitude of 32km. A small number of 4MJ/kg experiments were conducted at simulated flight altitudes of between 30 and 38km; the flight Mach number for these experiments was approximately 9.0. Thrust was calculated by integrating the centreline pressure distribution over the area of the thrust surface, assuming that the pressure at any axial location was constant across the engine width. These experimental thrust values were compared with theoretical estimates obtained using a one-dimensional analysis and a quasi-two-dimensional analysis. The comparison provided an indication of the level of completion of combustion in the experiments. The difference in thrust produced as a result of combusting fuel was examined by plotting the incremental specific impulse against equivalence ratio. Experimental and theoretical results agreed best at the higher equivalence ratios. Turbulent boundary layer separation correlations were used to provide reasonable estimates for the equivalence ratio at which the flow choked. The drag on the internal flowpath of the scramjet engine was estimated using the quasi-two-dimensional analysis. This drag estimate was combined with the experimental thrust measurements to provide estimates of the net specific impulse. Positive net specific impulse estimates were obtained above a certain minimum equivalence ratio, which depended on the contraction ratio and the test condition. The engine performance was observed to be highly dependent on the two-dimensional shock structure within the engine. Thrust and specific impulse were observed to decrease with increasing simulated flight altitude, as expected. Positive net specific impulse estimates were obtained at equivalence ratios of approximately one for simulated flight altitudes below 35km. Assuming complete combustion and that an equivalence ratio of one can be reached, the configuration considered in the present study can theoretically reach a net specific impulse of approximately 1000s at the 3MJ/kg condition and 500s at the 4MJ/kg condition. These numbers provide a promising testimonial for the use of this configuration, with modifications, as a more efficient alternative to rocket engines.

Combustion Instabilities In Solid Propellant Rocket Motors

Abstract: : These notes for two lectures are intended to provide the basic ideas for understanding and interpreting coherent oscillations is solid propellant rocket motors. The discussion is concerned mainly with the dynamics of a system consisting of two coupled sub-systems: the chamber containing combustion products; and the combustion processes con ned almost entirely to a thin region adjacent to the surface of burning propellant. Coupling between the sub-systems is always present due to the sensitivity of the ocmbustion processes to local values of pressure and velocity. Thus the primary mechanisms for instabilities in solid rockets are related to those interactions. A second mechanism involves vortex shedding, a cause of instabilities mainly in large motors, notably the Space Shuttle and Ariene V boost motors. Following a brief review of the history of combustion instabilities in solid rockets, the mechanisms and their quantitative representations are discussed. The remainder of the lectures is devoted to an approximate analysis providing a general framework convenient for understanding, predicting and interpreting combustion instabilities.

REDTOP-2: Rocket Engine Design Tool Featuring Engine Performance, Weight, Cost, and Reliability

Abstract: The Rocket Engine Design Tool for Optimal Performance-2 (REDTOP-2) is a newly created engineering design tool for use in the conc eptual and preliminary design of space transportation systems utilizing liquid propulsion rocket engines. REDTOP-2, one of many unique engineering tools commercially available from SpaceWorks Engineering, Inc. (SEI), represents a novel entry into the current suite of propulsion modeling tools. REDTOP-2 is capable of analyzing the flowpath characteristics of numerous engine configurations to perform a power balance of the turbomachinery hardware (pumps and turbines) to achieve a user specified main chamber combustion pressure. The engine performance, in terms of thrust and specific impulse (Isp), is then determin ed based on the results of this power balance and the flow conditions (pressure, temperat ure, flowrate, etc.) in the chamber(s) and nozzle(s). Engine weight is assessed at the main co mponent level using a combination of empirical and physics based analysis methods to provide vacuum, ambient, and sea-level thrust-to-weight (T/W) values. A cost model capable of predicting engine development, first unit, and production costs has been incorporated. Additionally, REDTOP-2 features a topdown modeling approach for computing engine safety and reliability metrics. REDTOP-2 is written in the modern, object-oriented C++ programming language and will execute on PC, Mac, and SGI platforms. Execution times are on the order of 30 seconds to 5 minutes, depending on the computing platform, engine configuration and design option selected by the user. User interface options currently include a command-line execution with ASCII file manipulation, filewrappers for use in Phoenix Integ ration’s ModelCenter© environment, and a PC-based graphical user interface (GUI). This paper will describe the REDTOP-2 tool and its capabilities. Sample results obtained from exercising the tool for a number of different existing engine designs will be presented . Results from a multi-variable sensitivity study on a LOX/LH2 fuel-rich, single preburner staged-combustion engine will be highlighted. Two sample applications involving vehi cle designs will be discussed. The first involves probabilistic/uncertainty analysis for an all-rocket vehicle design and the second the rocket main propulsion system analysis of an airbre athing, two-stage RLV concept with first stage tail-rockets and all-rocket second stage prop ulsion. Finally, future directions in the development of REDTOP-2 will be discussed.

Using Heterogeneous Propellant Burning Simulations as Subgrid Components of Rocket Simulations

Abstract: We discuss the manner in which one-dimensional unsteady descriptions can be constructed from multidimensional unsteady simulations of heterogeneous propellant combustion. Spatial averaging of the heat equation within the propellant is used to generate a one-dimensional equation with a number of source terms defined by the multidimensional thermal field and surface corrugations. Each of these terms is evaluated numerically, and those that can be neglected are identified; models are defined and tested for those that cannot. Closure of the one-dimensional description is achieved by relating the mean surface regression rate and the heat flux from the combustion field at the surface to the pressure and the average surface temperature. These relations are in the form of a look-up table generated from the “exact” (multidimensional) simulations. The accuracy of the one-dimensional system is tested by comparing the predictions with those of the exact model. This is done for steady burning rates at various pressures and for the unsteady burning response to pressure ramps and pressure pulses.

Feasibility and Performance of the Microwave Thermal Rocket Launcher

Abstract: Beamed‐energy launch concepts employing a microwave thermal thruster are feasible in principle, and microwave sources of sufficient power to launch tons into LEO already exist Microwave thermal thrusters operate on an analogous principle to nuclear thermal thrusters, which have experimentally demonstrated specific impulses exceeding 850 seconds Assuming such performance, simple application of the rocket equation suggests that payload fractions of 10% are possible for a single stage to orbit (SSTO) microwave thermal rocket We present an SSTO concept employing a scaled X‐33 aeroshell The flat aeroshell underside is covered by a thin‐layer microwave absorbent heat‐exchanger that forms part of the thruster During ascent, the heat‐exchanger faces the microwave beam A simple ascent trajectory analysis incorporating X‐33 aerodynamic data predicts a 10% payload fraction for a 1 ton craft of this type In contrast, the Saturn V had 3 non‐reusable stages and achieved a payload fraction of 4%

The Effects of Pressure and an Acoustic Field on a Cryogenic Coaxial Jet

Abstract: : A coaxial injector was designed to inject liquid nitrogen (LN2) with a coflow of gaseous nitrogen (GN2) in its annular region as part of a program to better understand the nature of the interaction between acoustic waves and liquid the jets in cryogenic rocket engines. Backlit images were taken from the jets at various flow rates and at sub-, near-, and super-critical chamber pressures with and without the presence of a 2700 Hz standing wave acoustic field. injector exit plane temperature measurements were made in both the center jet and annular regions. Results indicate that when the jet core appeared short and "thin", mostly under supercritical chamber pressures, the jet became insensitive to the external acoustic field. The strongest interaction was observed when the jet core looked long and "thick". To explore their implications, the characteristic acoustic impedance of the central jet and fuel/oxidizer momentum ratios are considered to play a role in the observed interactions. It is feasible that they play a similar role in cryogenic rocket engine combustion instability of the coaxial jet.

Gravity waves in the arctic mesosphere during the MaCWAVE/MIDAS summer rocket program

Abstract: [1] The MaCWAVE/MIDAS rocket campaign occurred at the Andoya Rocket Range (69°N,16°E) on July 1–2 and 4–5, 2002. This paper investigates gravity waves in the mesosphere using falling spheres dropped from rockets and the Weber sodium lidar at the ALOMAR observatory. The vertical displacement of a sodium sporadic layer on July 5 showed great variability at periods from minutes to hours with an observed frequency spectral slope of −1.89. The 2 salvos had similar wave amplitudes at the mesopause, whereas Salvo 2 had stronger amplitudes in the lower atmosphere. The dominant wave period varied strongly with height, possibly due to wave breaking on the strong mean gradients or oblique propagation of wave packets. One long-period wave appeared to propagate vertically from 75–95 km with a reduction of its vertical wavelength consistent with the mean wind gradient, but it is unclear whether it was a single wave or a superposition of waves.

Performance Prediction of a Ducted Rocket Combustor Using a Simulated Solid Fuel

Abstract: The ducted rocket is a supersonic flight propulsion system that takes the exhaust from a solid fuel gas generator, mixes it with air, and burns it to produce thrust. To develop such systems, the use of numerical models based on computational fluid dynamics (CFD) has been increasing, but to date only simplified treatments of the combustion within ducted rockets have been reported, likely due to the difficulties in characterizing and accurately modeling the partially reacted, particle-laden fuel exhaust from the gas generator. Through a careful examination of the governing equations and experimental measurements, a CFD-based methodology that properly accounts for the influence of the gas generator exhaust, particularly the solid phase, has now been developed to predict the performance of a ducted rocket combustor using a simulated solid fuel. It uses an equilibrium-chemistry probability density function combustion model with two separate streams, one gaseous and the other of 75-nm-diam carbon spheres, to represent the exhaust products from the gas generator. After extensive validation with direct-connect combustion experiments over a wide range of geometries and test conditions, this CFD-based method was able to predict, within a good degree of accuracy, the combustion efficiency of a ducted rocket combustor.

Three-axis attitude control propulsion device and flying object comprising the same

Abstract: A three-axis attitude control propulsion device and a flying object like a rocket including the device are provided in which combustion gas for attitude control can be efficiently used. A three-axis attitude control propulsion device 4 , having six nozzles has a motor case 6 and three-way discharge changeover valves 10, 10 ′ of a valve plug rotation type enabling a changeover of a flow passage by rotation of the valve plug. Combustion gas 18 is generated by combustion of propellant 8 in the motor case 6 . The three-axis attitude control propulsion device is operated so that one or two of the nozzles are opened to thereby discharge the combustion gas 18 and the remaining five or four nozzles are fully closed. Thereby, a three-axis attitude control of pitch control, roll control and yaw control, and control of a neutral state, can be selected.

Classification of Gas-Dust Formations from Rocket Exhaust in the Upper Atmosphere

Abstract: The results of research into the optical phenomena produced by rocket exhaust products in the upper atmosphere are presented. The data were obtained during routine observations of auroras by all-sky cameras from 1975 to 1990 from the Kola peninsula and Arckchangelsk region. The observed rocket launches were carried out from the Plesetsk and White Sea launch sites during both nighttime and twilight periods. The observed phenomena can be divided into two main types: local phenomena with long development times and relatively short-lived large-scale ones. The characteristic properties of both types are determined, in the first instance, by the type of rocket engines used (solid or liquid propellant) and their operating mode. The most intense, large-scale and dynamic phenomena are caused by separation of rocket stages and shutoff of solid-fuel rocket engines.

Maximum-likelihood rocket identifier

Abstract: A sensor suite determines location and velocity information relating to a missile threat, which is converted to missile or rocket state estimates. The state estimates are transformed into time-invariant dynamic parameters, unique for each missile type. Estimated rocket dynamic parameters are computed for each target type being considered, and compared with a reference set of rocket parameters representing different target types. The estimated rocket parameters are compared with the reference parameters in a maximum-likelihood sense, and combined using fuzzy logic to identify the rocket type and the likelihood. The identified rocket type and likelihood is used to aid in determining the future location of the missile so countermeasure can be applied.

A proposed bare-tether experiment on board a sounding rocket

Abstract: A mission on board a sounding rocket to carry out two bare-tether experiments is proposed: a test of orbital-motio n-limited (OML) collection and the proof-of-flight of a technique to determine the (neutral) density vertical profile in the critical E-layer. Since full bias from the motional field will be small (~ 20V), corresponding to a tape 1 km long and vrocket << 8 km/s, a power source with a range of supply voltages of few kV would be used. First, the negative terminal of the supply would be connected to the tape, and the positive terminal to a round, conductive boom of length 10 - 20 m; electrons collected by the boom cross the supply into the tape, where they leak out at the rate of ion impact plus secondary emission. Determination of the density profile from measurements of auroral emissions observed from the rocket, as secondaries racing down the magnetic field reach an E-layer footprint, are discussed. Next the positive terminal of the voltage supply is connected to the tape, and the negative terminal to a Hollow Cathode (HC); electrons now collected by the tape cross the supply, and are ejected at the HC. The opposite connections, with current collection operated by tape and boom, and operating on electrons and ions, and through partial switching in the supply, allow testing OML collection in almost all respects it depends on.

Numerical Analysis of Sloshing and Wave Breaking in a Small Vessel by CIP-LSM

Abstract: For the prediction of sloshing in the propellant tank of a rocket vehicle, the preliminary investigation was conducted. The flow field in the propellant tank during the ballistic flight of the vehicle was experimentally reproduced with the sub-scale model. The lateral acceleration as large as about 0.8G was provided with a mechanical exciter and the deformation of the liquid surface in the small vessel was visualized with a high-speed camera. The sloshing phenomena were also simulated with the CFD code, called CIP-LSM. The important features of surface deformation and wave breaking were successfully reproduced in the computation.

Classification of gas-dust formations from rocket exhaust in the upper atmosphere by scale and dynamics

Abstract: The results of research into the optical phenomena produced by rocket exhaust products in the upper atmosphere are presented. The data were obtained during routine observations of auroras by all-sky cameras from 1975 to 1990 from the Kola peninsula and Arckchangelsk region. The observed rocket launches were carried out from the Plesetsk and White Sea launch sites during both nighttime and twilight periods. The observed phenomena can be divided into two main types: local phenomena with long development times and relatively short-lived large-scale ones. The characteristic properties of both types are determined, in the first instance, by the type of rocket engines used (solid or liquid propellant) and their operating mode. The most intense, large-scale and dynamic phenomena are caused by separation of rocket stages and shutoff of solid-fuel rocket engines.

Resonant Motion of a Spin-Stabilized Thrusting Spacecraft

Abstract: The attitude instability of a spin-stabilized, thrusting upper stage spacecraft is investigated based on a two-body model consisting of a symmetric main body, representing the spacecraft, and a spherical pendulum, representing the liquefied slag pool entrapped in the aft section of the rocket motor. Exact time-varying nonlinear equations are derived and used to eliminate the drawbacks of conventional linear models. To study the stability of the spacecraft's attitude motion, both the spacecraft and pendulum are assumed to be in states of steady spin about the symmetry axis of the spacecraft and the coupled time-varying nonlinear equation of the pendulum is simplified. A quasi-stationary solution to that equation and approximate resonance conditions are determined in terms of the system parameters. The analysis shows that the pendulum is subject to a combination of parametric and external-type excitation by the main body and that energy from the excited pendulum is fed into the main body to develop the coning instability. When one of the resonance conditions and real flight data are used in the original time-varying nonlinear equations, the results match well with the observed motion before and after motor burnout of typical spin-stabilized upper stages. Some numerical examples are presented to explain the mechanism of the coning angle growth and how disturbance moments are generated.

Improvements to the Linear Standard Stability Prediction Program (SSP)

Abstract: Despite many decades of study, new solid rocket motors systems frequently experience unsteady gas motions and associated motor vibrations. This phenomenon most often occurs when the acoustic modes of the combustion chamber couple with combustion/flow processes. Current linear models of the sort used in the Standard Stability Prediction (SSP) code are designed to predict the tendency for a solid rocket motor to become unstable, but they do not provide any information on the severity of the instability (usually measured by the limit cycle amplitude of the oscillations) or on the triggerability (the tendency of an otherwise stable system to oscillate when pulsed with a sufficiently large disturbance) of the system. A goal of our present work is to build nonlinear capability into the the SSP tools. Success in incorporating useful nonlinear capabilities depends on a sufficiently complete and physically correct linear model. Accordingly, Software and Engineering Associates, Inc., has undertaken major improvements in the linear stability analysis and associated capabilities of the Solid Performance Program (SPP). These improvements support the development of new nonlinear capabilities to predict the oscillating pressure limit cycle amplitude, triggering and the DC pressure shift, the latter of which is often the most important threat to the rocket motor system resulting from combustion oscillations. In this paper, we focus on the recent additions and improvements to the linear SSP module. These include major improvement to the linkage between the rocket design code and SSP, and the inclusion of rotational flow effects that allow the satisfaction of key boundary conditions in the unsteady flow field solutions. The enhanced capability of the SSP is demonstrated by comparing the modified code to previous analyses for several solid rocket motors covering a wide range of typical design characteristics. Examples include systems predicted to be stable by earlier versions of the SSP code that were in fact inherently unstable. The improved linear code yields results which better fit the experimental findings.

Conceptual study of a combined cycle engine for an aerospace plane

Abstract: Operating conditions and performances of a fixed-geometry combined-cycle engine for a single-stage-to-orbit aerospace plane were calculated with a simple simulation model. With the flow conditions calculated with the model, the cooling requirement of the engine and pitching moment of the plane were investigated. The engine was composed of an ejector-jet mode, a ramjet mode, a scramjet mode, and a rocket mode. The engine had a fixed geometry in its operation. Subsonic combustion was conducted with no second throat in the combustor under the ejector-jet mode and the ramjet mode. Propellants were liquid hydrogen and liquid oxygen. The coolant flow rate became larger than the fuel flow rate. The excessive flow rate decreased the specific impulse above Mach 9 and restricted application of the airbreathing engine mode up to Mach 11. The pitching moment of the plane would be balanced even in the space in the configuration with the combined-cycle engine mounted on the windward surface.