Showing papers on "Rocket published in 2009"

Orbital Delivery of Small Payloads Using Hypersonic Airbreathing Propulsion

Abstract: Scramjet engines promise significantly higher specific impulse than rockets during the hypersonic phase of low- Earth-orbit insertion trajectories. Despite this, scramjets are not used on any current systems due to the difficulty of operating over the large Mach number envelope required by this accelerating trajectory. The key to taking advantage of airbreathing hypersonic engines for low-Earth-orbit insertion is to develop a multistage system that makes use of the scramjet only within its high-performance regime. Amultistage rocket-scramjet-rocket system that accepts this limitation has therefore been examined. This system includes a solid rocket boost to Mach 6, a near-term Mach 6–12 hydrogen-fueled scramjet engine to propel a reusable second stage, and a liquid-fueled final-stage rocket. Trajectory calculations for a system scaled to deliver approximately 100 kg to a 200 km equatorial orbit indicate payload mass fractions of approximately 1.5% with the use of a scramjet stage designed for low drag and efficient packaging. The goal of this work is to guide the future development of scramjets by identifying the areas that will make the most significant improvement to their use for space access.

Acoustics, Vortex Shedding, and Low-Frequency Dynamics Interaction in an Unstable Hybrid Rocket

Abstract: This paper deals with an experimental investigation into the stability behavior of a hybrid rocket where gaseous oxygen is fed with either an axial conical subsonic nozzle or a radial injector. The influence of the oxidizer-injection configurations on the motor stability is thoroughly examined. These distinct oxidizer-injection techniques allowed unveiling key and so far unreported features of the hybrid rocket combustion stability, especially emphasizing the role of vortex shedding which occurs in both the pre- and postcombustion chamber. Axial and radial injectors caused completely stable and unstable combustor operations, respectively, and this fact has been attributed to the fluid dynamics and unsteady heat release at the entrance of the fuel grain port. In particular, the unstable combustion in the radial-flow injector motor was dominated by low-frequency pressure oscillations, around 10-20 Hz. These low-frequency pressure oscillations were always accompanied by longitudinal acoustic modes. In some cases, the pressure oscillations abruptly increased, reaching peak-to-peak amplitude close to 70% of the mean chamber pressure, which is somewhat unusual for hybrid engines. Vortex shedding in the aft-mixing chamber is considered as the main driving mechanism of this latter behavior.

Coupled Analysis of Flow and Surface Ablation in Carbon-Carbon Rocket Nozzles

Abstract: A study is conducted to predict C/C nozzle recession behavior in solid rocket motors for broad variations of propellant formulations and motor operating conditions. The numerical model considers the turbulent flow in the nozzle, heterogeneous chemical reactions at the nozzle surface, variable transport and thermodynamic properties, and heat conduction in the nozzle material. Results show that the recession rate is largely determined by the diffusion of the major oxidizing species (H2O, CO2, OH) to the nozzle surface. Both the concentration of the major oxidizing species -affected by the aluminum content of the propellant- and the chamber pressure exert a strong influence on the recession rate. The erosion rate increases almost linearly with chamber pressure and decreases with propellants with higher aluminum content. The calculated results show a very good agreement with the experimental data from the BATES motor firings.

Limits on the Space Launch Market Related to Stratospheric Ozone Depletion

Abstract: Solid rocket motors (SRMs) and liquid rocket engines (LREs) deplete the global ozone layer in various capacities. We estimate global ozone depletion from rockets as a function of payload launch rate and relative mix of SRM and LRE rocket emissions. Currently, global rocket launches deplete the ozone layer ∼0.03%, an insignificant fraction of the depletion caused by other ozone depletion substances (ODSs). As the space industry grows and ODSs fade from the stratosphere, ozone depletion from rockets could become significant. This raises the possibility of regulation of space launch systems in the name of ozone protection. Large uncertainties in our understanding of ozone loss caused by rocket engines leave open the possibility that launch systems might be limited to as little as several tens of kilotons per year, comparable to the launch requirements of proposed space systems such as spaceplanes, space solar power, and space reflectors to mitigate climate change. The potential for limitations on launch syst...

Liquid-Propellant Rocket Engine Throttling: A Comprehensive Review

Abstract: Liquid-Propellant Rocket Engines (LREs) are capable of on-command variable thrust or thrust modulation, an operability advantage that has been studied intermittently since the late 1930s. Throttleable LREs can be used for planetary entry and descent, space rendezvous, orbital maneuvering including orientation and stabilization in space, and hovering and hazard avoidance during planetary landing. Other applications have included control of aircraft rocket engines, limiting of vehicle acceleration or velocity using retrograde rockets, and ballistic missile defense trajectory control. Throttleable LREs can also continuously follow the most economical thrust curve in a given situation, compared to discrete throttling changes over a few select operating points. The effects of variable thrust on the mechanics and dynamics of an LRE as well as difficulties and issues surrounding the throttling process are important aspects of throttling behavior. This review provides a detailed survey of LRE throttling centered around engines from the United States. Several LRE throttling methods are discussed, including high-pressure-drop systems, dual-injector manifolds, gas injection, multiple chambers, pulse modulation, throat throttling, movable injector components, and hydrodynamically dissipative injectors. Concerns and issues surrounding each method are examined, and the advantages and shortcomings compared.

Performance Validation of a Single-Tube Pulse Detonation Rocket System

Abstract: A pulse detonation engine (PDE) can be operated even if there are no compression mechanisms such as compressors or pistons, and a rocket engine with an extremely low combustor fill pressure (pulse detonation rocket, PDR) thus becomes possible. In this research, we made a model PDR system with increased specific impulse by partial fill. The performance predicted by this model was then confirmed experimentally. The thrust can be calculated by using the simplified PDE model of Endo et al. and the partial filling effect models of Sato et al. The mass flow rate of the propellant supplied from the pressurized cylinders is considered in this calculation. As a result, the thrust performance can be determined by the kind of propellant, the initial conditions of the gas in the cylinders, the supply-valve orifice and PDE-tube volume, and the operation frequencies. We fabricated a pulse detonation rocket (PDR) named “TODOROKI” and verified the thrust calculation model via a horizontal sliding test. We confirmed that the stability of the PDE operation depends on the ratio between the purge-gas thickness and the tube diameter. The thrust predicted by the model was identical to experimental results within 4%.

Film Cooling of Accelerated Flow in a Subscale Combustion Chamber

Abstract: Experimental investigations have been carried out to examine film cooling effectiveness of an accelerated hot gas in a subscale rocket combustion chamber. In support of future first-stage high-performance rocket combustion chambers, a Vulcain2-like test case has been examined with combustion pressure levels up to 12 MPa. The effectiveness of an almost tagentially injected film of hydrogen with an initial temperature of approximately 280 K has been determined. Axial distributions of temperature were measured inside the copper liner as well as on the chamber surface in the convergent and divergent parts of the nozzle segment. An existing film cooling model has been modified for application in a combined convective and filmcooled combustion chamber with an accelerated hot gas. The new model predicts film cooling effectiveness at different combustion-chamber pressures and film blowing rates at sub-, trans-, and supersonic conditions.

Sustained Hypersonic Flight in Europe: Technology Drivers for LAPCAT II

Abstract: Lapcat II is a logical follow-up of the previous EC-project which has as objective to reduce antipodal flights to less than 2 to 4 hours. Among the several studied vehicles, only two novel aircraft for a Mach 5 and 8 flight are retained in the present proposal. Starting from the available Mach 5 vehicle and its related precooled turbo-ramjet, assumed performance figures of different components will now be assessed in more detail. Though the cruise flight of the Mach 8 vehicle based on a scramjet seemed feasible, the acceleration based on an ejector rocket was not. Integrated design of airframe and engine throughout the whole trajectory is now the prime focus to seek for a conceptual design. A turbo-based engine will replace the former ejector rocket to assure better performance and fuel consumption during acceleration. Validated tool development should give solid confidence to propose a fully integrated vehicle to comply with the mission goals. Finally, for vehicles flying at high speeds and high altitudes, limited know-how is available on the environmental impact. The influence of NOx and H2O onto the ozone layer and the formation of contrails with its direct and indirect effects will be investigated for both vehicles.

Progress in Combustion and Heat Transfer Modelling in Rocket Thrust Chamber Applied Engineering

Abstract: The paper outlines the current status of Astrium's combustion and heat transfer prediction tools with special focus on latest model improvements. In particular heat transfer phenomena along overcooled walls including condensation, the modelling of closed cycle engine main combustion chamber injection conditions as well as oxygen/methane combustion and methane cooling prediction capabilities are addressed. First, the progress is manifested by dedicated test case simulations evaluating the extended capabilities by comparison to experimental data. Then, examples are given, how these model improvements contribute to master the challenges of future rocket thrust chambers developments.

Thermophysical Properties Measurements of Rocket Propellants RP-1 and RP-2

Abstract: The density, speed of sound, and viscosity of two rocket propellants (RP-1 and RP-2) have been measured. Densities were measured with two different instruments. Data at ambient atmospheric pressure were obtained with a rapid characterization instrument from 278.15 to 343.15 K that measured the speed of sound and density of the liquids in parallel. Adiabatic compressibilities derived from that data are included here. Densities of the compressed liquids were measured in an automated apparatus from 270 to 470 K and pressures to 40 MPa. Viscosities of the two liquids were measured in an open gravitational capillary viscometer at ambient atmospheric pressure from 293.15 to 373.15 K. The measurement results are consistent with compositional differences between the two samples. Correlations have been developed to represent the measured properties within the estimated uncertainties of the experimental data and to allow physically meaningful extrapolations beyond the range of the measurements.

Animal studies of the subgravity state during rocket flight

Abstract: I N T H E I R D I S C U S S I O N of man under gravity free conditions, Gauer and Haber G were the first to suggest that subgravity might lead to interesting physiological effects. They did not anticipate direct disturbances of respiration and of the circulation. However, they thought that the conflict between normal visual observations and the changed impulses from the labyrinth and proprioceptive apparatus might eventually lead to disturbances in performance and even to motion sickness. At the time the), proposed this, no practical means for producing the weightless state was available. Subsequently, Haber and Habeff pointed out that this condition could be briefly reproduced by an aircraft or rocket which followed a ballistic trajectory. I f the vehicle follows a course in which the horizontal velocity is constant and a downward acceleration exactly equaling the earth's gravity is developed, then no support is provided to objects within it and they become essentially weightless. Throughout the period that a high altitude research rocket coasts to its Presented at the Twenty-third Annual Meeting of the Aero Medical Association, Washington, D. C., March 17-19, 1952. This paper should be read in conjunction with those appearing in the Space Medicine Section in the August issue. Unfortunately, it was received too late to include with the other papers.

Energy-Based Acoustical Measurements of Rocket Noise

Abstract: Energy-based acoustical measurements are investigated in the context of more fully characterizing rocket noise source regions. Near-field measurements made on statically fired GEM-60 motors are described and the performance of two types of four-microphone tetrahedral probes is discussed. Vector intensity plots reveal the magnitude and directionality of the near-field sound radiation as a function of frequency, position, and time in the plume. HE development of the next-generation space flight vehicles has prompted renewed interest regarding source characterization and near-field propagation models of rocket noise. This source characterization is required to determine the vibroacoustic impact on flight hardware and structures in the vicinity of the launch pad. Brigham Young University has been involved in an effort to develop and validate an energy-based acoustic probe suitable for use in rocket fields, in particular the RS-68B engine to be used on the Ares V vehicle. Energy-based acoustical measurements require estimation of both the collocated acoustic pressure and the threedimensional particle velocity. From the pressure, a scalar, and the particle velocity vector, a number of energybased quantities can be calculated, including vector acoustic intensity, specific acoustic impedance, potential, kinetic, and total energy densities, and the Lagrangian density. Knowledge of one or more of these quantities may

Acoustic Measurement and Prediction of Solid Rockets in Static Firing Tests

Abstract: Acoustic measurements are executed in two series of static-firing tests of a solid rocket motor. The obtained data are quantitatively compared with calculation results of an empirical prediction method, NASA SP-8072 and CFD. According to the results, the NASA SP-8072 overestimates the sound pressure levels at the 20° and 35° points from the jet axis in the far field, although the SPLs at other measured points are reasonably predicted. On the other hand, the CFD calculation can clearly explain the generation and propagation mechanism of the acoustic wave and reasonably predict the SPLs at all the measured points. From the results, it is confirmed that the prediction accuracy of the CFD calculation is within 5 [dB] in overall sound pressure level, which is within the experimental uncertainty involved in the measured data, and the CFD is effective for the prediction of both the near and the far field acoustics generated from the rocket motors.

Aluminum-Ice (ALICE) propellants for hydrogen generation and propulsion

Abstract: An experimental investigation was conducted to determine the relative propulsive performance and viability of novel solid propellants comprised of ALICE using fundamental techniques such as steady-state strand experiments and applied experimentation such as labscale static fire rocket tests. Burning rates, slag accumulation, thrust, and pressure are some of the experimental parameters obtained. System scaling has been performed to examine the effect of larger systems on slag accumulation and performance parameters. The effect of pressure on the linear burning rate was examined and correlated using a Saint Roberts’s law fit. The pressure exponent for ALICE was 0.73, which is approximately a factor of two larger than Al/water mixtures. Three sizes rocket motors ranging from internal diameters of 0.75 to 3-in. Nozzle throat diameter and igniter strength were varied. It was found that ALICE propellants successfully ignited and combusted in each lab-scale rocket motor, generating thrust levels above 223 lbf for expansion ratios of 10 and center-perforated grain configurations (3-in length). For the 3-in motor, combustion efficiency was around 70%, while the specific impulse efficiency was 64%.

Development and Launch of the Purdue Hybrid Rocket Technology Demonstrator

Abstract: The Purdue University School of Aeronautics and Astronautics is developing a hybrid rocket technology demonstrator to serve as a test bed for technologie s critical to the development of vehicles capable o f delivering microgravity experiments to altitudes ex ceeding 100 km. These critical technologies include propulsion, structures, separation, recovery, groun d support, avionics and guidance, navigation and co ntrol sub-systems. These technologies will be demonstrated sequentially over a series of test flights which will allow the designers to validate each of these sub-systems before adding more complexity, risk and features t o the technology demonstrator. To date, three different h ybrid rocket motors have been designed, manufactured and tested. In addition the Ground Support Equipment required for transferring the oxidizer to and fro m the flight-vehicle has been constructed and tested exte nsively during cold flow and hot-fire test operatio ns. A Mobile Launch Platform was also developed for transporting the ground support equipment and for performing launch operations in designated areas. This paper details the design and development of a f lightweight, 900 lbf thrust, 90% hydrogen peroxide/LDPE hybrid rocket motor which was successfully hot-fire tested in vertical configuration a total of five ti mes at the Purdue Vertical Rocket Test Facility. In addition the paper describes the successful launch of the 1s t generation hybrid flight-vehicle which reached an altitude of 6,100 ft (Mach 0.6) in June 2009, making a first important step towards flight operations for this series of hybrid rocket technology demonstrators.

Technological Aspects of Transpiration Cooled Composite Structures for Thrust Chamber Applications

Abstract: The long-term development of transpiration cooled ceramic rocket thrust chambers at the German Aerospace Center (DLR) currently culminates in designs of self-sustaining fibre reinforced rocket engine chamber structures. This paper explains characteristic issues and potential benefits introduced by this new technology, which seem to be achievable in terms of weight and cost reduction, increased reliability and higher lifetime due to no longer existing thermal cycling sensitivity. The paper furthermore describes design and functional aspects of the chamber, the component manufacturing process and shows some experimental results of test campaigns with respect to structure and material relevance. DLR’s development road map is sketched and the technology readiness level achieved so far is discussed. Nomenclature ISP = specific impulse kd = coefficient of Darcyan permeability kf = coefficient of non-Darcyan (Forchheimer) permeability L = flow length pin = inflow pressure pout = outflow pressure T = temperature v = velocity Δp = pressure loss λ = thermal conductivity (CTC) e = porosity η = dynamic viscosity ρ = density

Investigation of Two Dimensional Thermal Loads in the Region Near the Injector Head of a High Pressure Subscale Combustion Chamber

Abstract: Optimization of heat transfer management is a key issue in designing a rocket combustion chamber. This paper presents the new test specimen and the measurement technique that has been developed und successfully used by the Institute of Space Propulsion for high resolution investigations of the influence of injector-wall interactions on heat transfer in a subscale combustion chamber. The new measurement method allows obtaining two dimensional heat load distribution on the hot gas side at real rocket engine-like conditions at combustion chamber pressures up to 15 MPa. The presented investigations have been performed at the European Research and Technology Test Facility P8 for cryogenic subscale rocket engines.

ARTIFICIAL AGEING OF DOUBLE BASE ROCKET PROPELLANT Effect on dynamic mechanical properties

Abstract: The ageing of double base rocket propellants (DB rocket propellants), which is a consequence of chemical reactions and physical processes that take place over time, has significant effect on their relevant properties (e.g. chemical composition, mechanical properties, ballistic properties, etc.). The changes of relevant properties limit the safe and reliable service life of DB rocket propellants. This is the reason why numerous research efforts are devoted to finding out reliable methods to measure the changes caused by ageing, to assess the quality at a given moment of time, and to predict remaining life-time of DB rocket propellants. In this work we studied dynamic mechanical properties of DB rocket propellant artificially aged at elevated temperatures, in order to detect and quantify changes in dynamic mechanical properties caused by the ageing. Dynamic mechanical properties were studied using dynamic mechanical analyser (DMA). The results obtained have shown that the ageing causes significant changes of DMA curve’s shape and positions. These changes are quantified by following some characteristic points on DMA curves (e.g. glass transition temperatures; storage modulus, loss modulus and tanδ at characteristic temperatures, etc.). It has been found out that the most sensitive parameters to the ageing process are: storage modulus at viscoelastic and softening region, peak width and height on loss modulus curve, glass transition and softening temperature, and tanδ at viscoelastic region.

A Study of Nonlinear Combustion Instability

Abstract: Combustion instability (CI) has been persistent in all forms of propulsion since their inception. CI is characterized by pressure oscillations within the propulsion system. If even a small fraction of the dense energy within the system is converted to acoustic oscillations the system vibrations can be devastating. The coupling of combustion and fluid dynamic phenomena in a nonlinear system poses CI as a significant engineering challenge. Drawing from previous analysis, second order acoustic energy models are taken to third order. Second order analysis predicts exponential growth. The addition of the third order terms capture the nonlinear acoustic phenomena (such as wave steepening) observed in experiments. The analytical framework is derived such that the energy sources and sinks are properly accounted for. The resulting third order solution is compared against a newly performed simplified acoustic closed tube experiment. This experiment provides the interesting result that in a forced system, as the 2nd harmonic is driven, no energy is transferred back into the 1st mode. The subsequent steepened waveform is a summation of 2nd mode harmonics (2, 4, 6, 8...) where all odd modes are nonexistent. The current third order acoustic model recreates the physics as seen in the experiment. Numerical experiments show the sensitivity of the pressure wave limit cycle amplitude to the second order growth rate, highlighting the importance of correctly calculating the growth rates. The sensitivity of the solution to the third order parameter is shown as well. Exponential growth is found if the third order parameter is removed, and increased nonlinear behavior is found if it retained and as it is increased. The solutions sensitivity to this term highlights its importance and shows the need for continued analysis via increasing the models

Transient Combustion Response of AP/HTPB Composite Propellant to Acoustic Oscillations in a Rocket Motor

Abstract: A comprehensive theoretical/numerical model is developed to investigate the transient combustion response of AP/HTPB composite propellant to acoustic excitation in a rocket-motor environment. The work extends our previous analysis of AP/HTPB combustion at steady-state to include flow oscillations and their subsequent influence on the flame structure and propellant burning behavior. Detailed information about the flame-zone physiochemistry near the propellant surface is obtained at different locations in the motor for the first three modes of longitudinal acoustic waves. In addition, various mechanisms dictating the characteristics of the propellant combustion response, including microscale motions in the flame zone and macroscale motions in the bulk flow, are explored. The effects of mean and oscillatory flowfields in determining the propellant combustion response are also examined. Furthermore, a large flow velocity fluctuation often leads to a nonlinear response of the heat feedback to the propellant su...

Status Update Report for the Peregrine Sounding Rocket Project: Part III

Abstract: The Peregrine Sounding Rocket Program is a joint program of NASA-Ames, NASAWallops, Stanford University and Space Propulsion Group, Inc. to develop and fly a high performance sounding rocket based on liquefying fuel hybrid rocket technology. The program was kicked off in November of 2006 and initial ground testing of the propulsion system begain in July 2008. Two virtually identical vehicles capable of lofting a 5kg payload to 100km will be constructed and flown out of the NASA Sounding Rocket Facility at Wallops Island. The propellants utilized are nitrous oxide and paraffin, a high regression rate liquefying fuel initially developed at Stanford University. The goal of the Peregrine program is to demonstrate the operational maturity of liquefying hybrid propulsion systems for space applications and their potential to reduce propulsion system costs. This is the third in a series of three annual papers outlining the Peregrine project and providing status updates. The majority of this (JPC 2009) paper will focus on the results of the propulsion system ground test program and the detailed design of the vehicle.

SRM Internal Ballistic Numerical Simulation by SPINBALL Model

Abstract: In the design and development of solid propellant rocket motors (SRMs), the use of numerical tools able to simulate, predict and reconstruct the behavior of a given motor, in all its operative conditions, is particularly important in order to decrease all the planning times and costs. This paper is devoted to propose and present an approach to the numerical simulation of SRM internal ballistics, during the entire combustion time, by means of dierent own made models. The core of this procedure is represented by the SPINBALL model and numerical code. SPINBALL considers a Q1D unsteady modeling of the SRM internal ballistics, with many dierent sub-models able to represents all the driving phenomena that characterize the bore chamber oweld conditions during the SRM timelife, from the motor start-up to burn-out. In particular, the grain burning surface evolution is accomplished by means of a 3D numerical grain regression model, named GREG. This model is based on a full matrix level set approach, on rectangular or cylindrical structured grids. GREG gives to the SPINBALL gasdynamical model the evolution in time of the port area, wet perimeter and burn perimeter along the motor axis and, in case, within the submergence zone. The nal objective is, hence, to develop an analysis/simulation capability of SRM internal ballistics, for the entire combustion time, with simplied physical models, in order to reduce the computational cost required, but ensuring, in the meanwhile, an accuracy of the simulation greater than the one usually given by 0D quasi steady models, during quasi steady state and tail o. Notwithstanding, a 0D quasi steady model of SRM internal ballistic has been developed to reconstruct the experimental data coming from static ring tests (SFTs), in order to evaluate non-ideal behaviour parameters, like combustion eciency, hump law and nozzle eciency and the nozzle throat area evolution. These parameters are used in the SPINBALL model as inputs. The results of the internal ballistics numerical simulation, from motor start-up to burnout yielded with the SPINBALL model, will be shown for Zero23, second solid rocket motor stage developed in the ESA (European Space Agency) project of the new European small launcher Vega.

Automation of Structured Overset Mesh Generation for Rocket Geometries

Abstract: Aerodynamic characterization in the ascent phase is one of the necessary steps to designing a successful rocket. Computational fluid dynamics (CFD) simulations that employ the structured overset grid philosophy are commonly used for aerodynamics analysis. While the method is highly desirable due to its ability to provide viscous flow solutions for complex geometries, the preliminary geometry processing work required to generate grids is a major bottleneck to efficiently obtaining fluid simulation results. The present paper proposes strategies to improve the grid generation process by eliminating the mundane tasks that can be automated with limited knowledge of the geometry and little user input. The automation is targeted for rocket bodies and the protuberances that are commonly placed on rockets. However, the resulting tools may be applicable to other geometries. The present paper’s focus is on a scripting framework to automate surface and volume mesh generation from a native computer aided design (CAD) solid model geometry definition. The resulting process is able to generate overset meshes with fewer lines of code and less user input culminating in savings in time to process a clean solid model CAD geometry to a CFD-ready mesh.

Discrimination for ablative control mechanism in solid-propellant rocket nozzle

Abstract: The ablation in solid-propellant rocket nozzle is a coupling process resulted by chemistry, heat and mass transfer. Based on the heat and mass transfer theory, the aero-thermo-dynamic, and thermo-chemical kinetics, the thermal-chemical ablation model is established. Simulations are completed on the heat flow field and chemical ablation in the nozzle with different concentrations, frequency factors and activation energy of H2. The calculation results show that the concentration and the activation energy of H2 can provoke the transformation of control mechanism, whereas the influence brought by the frequency factor of H2 is feeble under a high-temperature and high-pressure combustion circumstance. The discrimination for ablative control mechanism is dependent on both concentration and activation energy of H2. This study will be useful in handling ablation and thermal protection problem in the design of solid-propellant rocket.

Experiment and Semi-Empirical Modeling of Lab-Scale Hybrid Rocket Performance

Abstract: A simplified semi-empirical predictive model was developed to aid in the determination of operating parameters and chamber specifications for a lab-scale hybrid rocket engine and test sled design. The model combines user defined initial operating and system design parameters with empirically derived regression rate correlations, NASA CEA2000 combustion equilibrium analysis results, and conservation of mass derivations. The model facilitates parametric optimization of oxidizer flow, chamber pressure and nozzle throat diameter, through a time resolved series of functions, deriving output parameters including characteristic velocity, combustion temperature, efficiency, chamber pressure, thrust, and specific inertia. Experiments were conducted using polymethyl methacrylate (PMMA), hydroxyl-terminated polybutadiene (HTPB) and gaseous oxygen. Experimental results indicates HTPB regression rate exceeds PMMA by a factor of 2 for a given oxidizer flow rate and nozzle parameters. Additionally, the results show, a simplified model of the hybrid combustion system is sufficient to adequately predict combustion parameters in a lab-scale hybrid rocket motor.