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Karabeyoglu A.,Space Propulsion Group, Inc | Stevens J.,Space Propulsion Group, Inc | Geyzel D.,Space Propulsion Group, Inc | Cantwell B.,Space Propulsion Group, Inc | Micheletti D.,MADA
47th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit 2011 | Year: 2011

Hybrid rocket propulsion is a tipping point technology in the sense that a small, short term investment could have game changing consequences in developing green, safe, affordable and high performance systems needed in future space missions. In order to demonstrate the advantages of hybrids most effectively, the effort should be concentrated on improving the Technology Readiness Level (TRL) of the technology for a carefully selected class of missions. Arguably upper stage motors used in small launch vehicles constitute a perfect platform for this purpose due their relatively small scale and high performance requirements. The advanced hybrid rockets that are being developed by SPG are believed to have the capability to deliver high performances desirable for upper stages, while retaining the cost, environmental and simplicity advantages of the classical hybrids. In order to demonstrate the performance capabilities of advanced hybrid rockets, a design study has been conducted to replace the Orion 38 solid rocket motor with a LOX/paraffin-based system. The LOX hybrids delivering the same level of total impulse as Orion 38 system are determined to be 15-18% lighter. It has been shown that switching to higher performance hybrid upper stages could lead to payload increases up to 40% for a typical launch vehicle. The additional cost, environmental, safety, stop/restart/throttling advantages are expected to make the hybrids desirable alternatives to the existing upper stage systems. © 2011 by Arif Karabeyoglu.


Micheletti D.A.,Universal Technical Resource Services, Inc. | Karabeyoglu,Space Propulsion Group, Inc
Proceedings of the International Astronautical Congress, IAC | Year: 2012

The classical hybrid rocket systems developed to date have suffered from two major shortcomings: 1) complex multiport fuel grains as a result of the poor regression rate performance of the classical polymeric fuels; and 2) low frequency instabilities. In the past, the mitigation methods for these problem areas have introduced significant complexity into the motor design, compromising the simplicity advantage of hybrid rockets. For example, the 250 klb motor developed by American Rocket Company (AMROC) was based on a complex 15 port wagon wheel configuration (resulting in poor fuel utilization and expensive fabrication), and the motor stability was achieved by the continuous injection of a hazardous pyrophoric substancc. triethylaluminum (TEA). Space Propulsion Group, Inc.'s (SPG's) paraffin-based/LOX hybrid rocket technology, which has an inherently high fuel regression rate, allows for the use of a simple single circular port fuel grain design approach. SPG has also developed unique proprietary technologies to eliminate the low frequency instabilities in LOX-based hybrids without resorting to external heat or pyrophoric liquid addition at the front end of the motor. These two technological advancements are crucial in keeping the hybrid concept cost effective, simple, and safe compared to the state-of-the art of liquid and solid rocket systems. Recently, SPG has successfully converted its 11-inch diameter heavy weight motor to a 10-inch carbon composite based flight weight system that can deliver 7,000 lb of vacuum thrust for 15 seconds. SPG has also started a scaled-up development effort of its paraffin-based/LOX motor technology. The 22-inch diameter scalcd-up flight weight motor is capable of producing thrust forces up to 35,000 lbf. Ground testing of both the 10-inch and 22-inch flight weight motors is currently ongoing at the Montana Aerospace Development Association (MADA) Butte AeroTec Facility, which is located near Butte, Montana in the United States. To date, eighteen flight weight motors have been successfully tested in the development program. Copyright © (2012) by the International Astronautical Federation.


Nozari H.,Koç University | Karabeyolu A.,Koç University | Karabeyolu A.,Space Propulsion Group, Inc
Fuel | Year: 2015

Abstract With its high hydrogen density and already existing infrastructure, ammonia (NH3) is believed to be an excellent green fuel that can be used in energy generation and transportation systems. Combustion of ammonia has certain challenges (associated with its low flame speed and fuel bond NOx emissions) that need to be addressed before its widespread use in practical systems. The primary objective of this study is to develop a reduced reaction mechanism for the combustion of ammonia which can be used to expedite the design of effective ammonia combustors through numerical simulations of realistic combustor geometries with accurate kinetics models. First we have investigated the combustion characteristics of NH3/H2/air mixtures at elevated pressure and lean conditions which are encountered in practical systems such as gas turbine combustors. Laminar premixed freely propagating flame model is used to calculate the combustion properties. The results of sensitivity study of total NOx formation with respect to the equivalence ratio indicates the possibility of localized rich combustion as an effective way to reduce the NOx concentration down to levels that are the same order as the modern gas turbine engines. In the second part of the study, by considering a wide range of conditions in terms of pressure, fuel mixture, and equivalence ratio we have developed two reduced mechanisms based on the Konnov mechanism. The reduced mechanisms are capable of predicting the total NOx emission level and the laminar flame speed at an acceptable accuracy over a wide range of conditions. Evaluating the performance of the reduced mechanisms with respect to the full mechanism and experimental data shows that the mechanisms are able to predict the combustion properties almost at the same accuracy level as the Konnov mechanism, but at a nearly five times less CPU time expense. © 2015 Elsevier Ltd.


Nozari H.,Koç University | Karabeyoglu A.,Space Propulsion Group, Inc
13th International Energy Conversion Engineering Conference | Year: 2015

In the first section of this numerical study we investigate the combustion characteristics of ammonia-air mixtures at elevated pressure and lean conditions which are encountered in gas turbine combustors. Laminar premixed freely propagating flame and homogenous reactor models are used to calculate the combustion properties. The improvement by hydrogen addition to the fuel mixture in combustion characteristics such as laminar flame speed and ignition delay time is noticeable. Based on ammonia decomposition sensitivity analysis, it is found that the OH radicals have a leading role in controlling the fuel mole conversion and the laminar flame speed. The results of sensitivity study of total NOx formation with respect to the equivalence ratio reveal the possibility of localized rich combustion as an effective way to reduce the NOx concentration down to levels that are the same order as the modern gas turbine engines. In the second part of the study, by considering a wide range of conditions in terms of pressure, fuel mixture, and equivalence ratio we develop two reduced mechanisms based on the Konnov mechanism. The reduced mechanisms are capable of predicting total NOx emission level and laminar flame speed in an acceptable accuracy under wide range of conditions. Evaluating performance of the reduced mechanisms with respect to the full mechanism and experimental data shows that the mechanisms are able to predict the combustion properties with almost the same accuracy as the full Konnov mechanism and with nearly five times less CPU time expense.


Chandler A.A.,Stanford University | Cantwell B.J.,Stanford University | Scott Hubbard G.,Stanford University | Karabeyoglu A.,Space Propulsion Group, Inc
Acta Astronautica | Year: 2011

Developments in hybrid propellants over the last decade make the hybrid motor a viable candidate and possibly an enabling technology for a Mars Ascent Vehicle (MAV) as part of a Mars Sample Return (MSR) campaign. Fast regression rate fuels such as paraffin allow for single port hybrid designs that overcome the disadvantages associated with classical hybrid fuels. Additionally, paraffin has a both a weak and low glass transition temperature, making it an ideal candidate for a Mars application. Nytrox, a high performance oxidizer made of a mixture of nitrous oxide and oxygen, can be chosen to match the temperatures on Mars. The hybrid MAV can survive the harsh Martian environment with minimal or no thermal conditioning and it can meet the design challenges posed by coordinating with the other aspects of MSR. © 2011 Elsevier Ltd. All rights reserved.


Kara O.,Koç University | Karabeyoglu A.,Koç University | Karabeyoglu A.,Space Propulsion Group, Inc
AIAA SPACE 2015 Conference and Exposition | Year: 2015

This study underlies small satellite architecture optimization by using existing electric propulsion systems for the Moon missions. The estimated objective is panoramic imaging of the Moon accompanied with future in-situ applications. Edelbaum’s low thrust trajectory transfer with optimal control theory is used to calculate the required ΔV. During the journey, 1.5h eclipse duration effects the solar array design. The optimized xenon propellant density and pressure are 1350 kgm3 and 8.3 MPa within 300K. Two types of optimization process revealed based on hexagonal SC architecture. The iterative method with LEO departured ion thruster has 23 mN with minimum 213 kg total mass. Corresponding SC volume is 0.70 m3, propellant mass is 64 kg. This scenario cost $108.5M and takes 980 days. Same thruster level for GEO departure case takes 880 days with 58 kg xenon gas. The total cost reduces $2.5M. For HALL engine design, LEO departure case needs 0.8 m3, 247 kg SC including 82 kg xenon. 77 mN thrust operates 208 days towards the Moon that ends up with $121M total cost. If the SC to be launched from GEO, flight time reduces 45 days by consuming 65 kg propellant. Total SC mass, volume and power values are 230 kg, 0.71 m3 and 1351W which cost $115M. Results are compared with previous Moon or electric propulsion missions such as SMART-1, LADEE, Clementine and Hayabusa. For future applications of small satellites, innovative concepts are envisioned for in-space, Earth-independent exploration and space education. © 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. All Rights Reserved.


Grant
Agency: Department of Defense | Branch: Navy | Program: SBIR | Phase: Phase I | Award Amount: 79.97K | Year: 2014

The use of a ramjet motor in a tactical system offers significant performance advantages over a typical boost sustain propulsion system. SPG, Inc. proposes to conduct a study demonstrating the feasibility of solid-fuel ramjet (SFRJ) utilizing paraffin-based solid fuels for tactical applications that require high speed and long range performance. In addition, a novel liquefying ramjet will be designed, aiming for improved combustion efficiency and flame holding characteristics. SFRJ propulsion offers many advantages over a traditional boost-sustain tactical system. Potential increases in range of more than 4 times that of a typical boost-sustain tactical system are possible using paraffin solid fuels in a ramjet configuration. The incorporation of energetic additives (e.g., boron) further increases the potential range of the system. Additives will be evaluated to determine the propulsive performance gains from inclusion. Selected fuel formulations will be evaluated using an in-house flyout code to calculate the theoretical range of the system. Fuel formulations for evaluation in Phase II will be selected at the end of the period. Static test firings will be completed with a baseline paraffin fuel. Tests will demonstrate the range of throttling and combustion efficiency will be measured for the baseline fuel in traditional SFRJ configuration.


Grant
Agency: Department of Defense | Branch: Navy | Program: SBIR | Phase: Phase I | Award Amount: 79.95K | Year: 2011

Space Propulsion Group, Inc (SPG) proposes to conduct investigations on ignition and barrier systems for multi-pulse rocket motors. The benefits of the proposed work is the reliable ignition of propellant grains under various free chamber volume conditions and the ability to terminate thrust generation for a desired amount of time following discrete propellant grain burnout. Capability to operate with multiple thrust pulses without the complexity of staging is beneficial for reducing propulsion system costs and increasing mission flexibility. The ability to terminate propulsive thrust following propellant grain combustion relies on the ability to physically and thermally isolate the discrete propellant grains. The significantly different chamber conditions at the start of each of the individual pulses require a separate approach for ignition. An ignition system that can be tailored to the initial conditions of each propellant grain allows great flexibility in selecting mission profiles. The Phase I work considers the development of the igniter system for discrete pulses and a pulse separation device. Evaluation and design of potential igniter systems as well as design and preliminary testing of the pulse separation device will be conducted.


Grant
Agency: National Aeronautics and Space Administration | Branch: | Program: SBIR | Phase: Phase I | Award Amount: 123.75K | Year: 2016

A Mars Sample Return (MSR) campaign has been identified as the critical next step in Mars science. Of the tasks needing to be accomplished, the Mars Ascent Vehicle (MAV), or means for getting the samples into orbit around Mars, is considered the highest risk. The MAV will be required to remain on the Martian surface for a year or more in order to return to Earth on a minimum energy trajectory and coordinate with the other aspects of the MSR. Environmental conditions on Mars are a significant concern, with seasonal extremes of about ?110?C and 25?C. To reduce the required system mass and power related to thermal management, fuel, oxidizer, and ignition system components should be able to withstand these temperature variations. The trajectory of the MAV also has the requirement for restart capabilities. The focus of this proposal is the development of hypergolic ignition systems for MAV application. The current oxidizer of choice for the MAV is a mixed oxides of nitrogen (MON), which is a known hypergol with many fuels. The proposed solution uses metal particles (e.g., hydrides, boranes, or borohydrides) to generate ignition by injection and mixing with the oxidizer. The particles are initially housed in a sealed, pressurized chamber prior to injection. An inert gas acts as the pressurant for the particles and also serves to protect the particles from oxidation and hydrolysis during storage. Elimination of polymeric materials from the ignition train eliminates concerns of glass transition, which could lead to ignition failure in traditional pyrotechnic/pyrogen igniters. The focus of the study is the design, analysis, and testing of potential particle injection configurations and the selection of the ideal candidate particle type and morphology. At the completion of Phase I, as system ready for subscale testing in a hybrid rocket motor will be ready for implementation.


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