46 CFR 160.040-4 - Equipment for impulse-projected rocket type line-throwing appliance.

Code of Federal Regulations, 2014 CFR

2014-10-01

... 46 Shipping 6 2014-10-01 2014-10-01 false Equipment for impulse-projected rocket type line...-Throwing Appliance, Impulse-Projected Rocket Type (and Equipment) § 160.040-4 Equipment for impulse-projected rocket type line-throwing appliance. (a) Four rocket projectiles, each complete with bridle and...

46 CFR 160.040-4 - Equipment for impulse-projected rocket type line-throwing appliance.

Code of Federal Regulations, 2010 CFR

2010-10-01

... 46 Shipping 6 2010-10-01 2010-10-01 false Equipment for impulse-projected rocket type line...-Throwing Appliance, Impulse-Projected Rocket Type (and Equipment) § 160.040-4 Equipment for impulse-projected rocket type line-throwing appliance. (a) Four rocket projectiles, each complete with bridle and...

46 CFR 160.040-4 - Equipment for impulse-projected rocket type line-throwing appliance.

Code of Federal Regulations, 2011 CFR

2011-10-01

... 46 Shipping 6 2011-10-01 2011-10-01 false Equipment for impulse-projected rocket type line...-Throwing Appliance, Impulse-Projected Rocket Type (and Equipment) § 160.040-4 Equipment for impulse-projected rocket type line-throwing appliance. (a) Four rocket projectiles, each complete with bridle and...

46 CFR 160.040-6 - Marking and labeling.

Code of Federal Regulations, 2013 CFR

2013-10-01

...: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Line-Throwing Appliance, Impulse-Projected Rocket Type (and... rocket-projectiles shall be legibly marked with the name of the manufacturer, the model designation, the...

46 CFR 160.040-6 - Marking and labeling.

Code of Federal Regulations, 2011 CFR

2011-10-01

...: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Line-Throwing Appliance, Impulse-Projected Rocket Type (and... rocket-projectiles shall be legibly marked with the name of the manufacturer, the model designation, the...

46 CFR 160.040-6 - Marking and labeling.

Code of Federal Regulations, 2014 CFR

2014-10-01

...: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Line-Throwing Appliance, Impulse-Projected Rocket Type (and... rocket-projectiles shall be legibly marked with the name of the manufacturer, the model designation, the...

46 CFR 160.040-6 - Marking and labeling.

Code of Federal Regulations, 2012 CFR

2012-10-01

...: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Line-Throwing Appliance, Impulse-Projected Rocket Type (and... rocket-projectiles shall be legibly marked with the name of the manufacturer, the model designation, the...

46 CFR 160.040-6 - Marking and labeling.

Code of Federal Regulations, 2010 CFR

2010-10-01

...: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Line-Throwing Appliance, Impulse-Projected Rocket Type (and... rocket-projectiles shall be legibly marked with the name of the manufacturer, the model designation, the...

Hybrid rocket propulsion

NASA Technical Reports Server (NTRS)

Holzman, Allen L.

1993-01-01

Topics addressed are: (1) comparison of the theoretical impulses; (2) comparison of the density-specific impulses; (3) general propulsion system features comparison; (4) hybrid systems, booster applications; and (5) hybrid systems, upper stage propulsion applications.

Specific Impulse and Mass Flow Rate Error

NASA Technical Reports Server (NTRS)

Gregory, Don A.

2005-01-01

Specific impulse is defined in words in many ways. Very early in any text on rocket propulsion a phrase similar to .specific impulse is the thrust force per unit propellant weight flow per second. will be found.(2) It is only after seeing the mathematics written down does the definition mean something physically to scientists and engineers responsible for either measuring it or using someone.s value for it.

46 CFR 160.040-7 - Procedure for approval.

Code of Federal Regulations, 2014 CFR

2014-10-01

...: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Line-Throwing Appliance, Impulse-Projected Rocket Type (and Equipment) § 160.040-7 Procedure for approval. (a) Rocket type line-throwing appliances are approved by the...

46 CFR 160.040-7 - Procedure for approval.

Code of Federal Regulations, 2011 CFR

2011-10-01

...: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Line-Throwing Appliance, Impulse-Projected Rocket Type (and Equipment) § 160.040-7 Procedure for approval. (a) Rocket type line-throwing appliances are approved by the...

46 CFR 160.040-7 - Procedure for approval.

Code of Federal Regulations, 2013 CFR

2013-10-01

...: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Line-Throwing Appliance, Impulse-Projected Rocket Type (and Equipment) § 160.040-7 Procedure for approval. (a) Rocket type line-throwing appliances are approved by the...

46 CFR 160.040-7 - Procedure for approval.

Code of Federal Regulations, 2012 CFR

2012-10-01

...: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Line-Throwing Appliance, Impulse-Projected Rocket Type (and Equipment) § 160.040-7 Procedure for approval. (a) Rocket type line-throwing appliances are approved by the...

46 CFR 160.040-7 - Procedure for approval.

Code of Federal Regulations, 2010 CFR

2010-10-01

...: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Line-Throwing Appliance, Impulse-Projected Rocket Type (and Equipment) § 160.040-7 Procedure for approval. (a) Rocket type line-throwing appliances are approved by the...

Combustion Performance of a Staged Hybrid Rocket with Boron addition

NASA Astrophysics Data System (ADS)

Lee, D.; Lee, C.

2018-04-01

In this paper, the effect of boron on overall system specific impulse was investigated. Additionally, a series of combustion tests was carried out to analyze and evaluate the effect of boron addition on O/F variation and radial temperature profiles. To maintain the hybrid rocket engine advantages, upper limit of boron contents in solid fuel was set to be 10 wt%. The results also suggested that, when adding boron to solid fuel, it helped to provide more uniform radial temperature distribution and also to increase specific impulse by 3.2%.

The Future of the U.S. Intercontinental Ballistic Missile Force

DTIC Science & Technology

2014-01-01

42 3.7. Nevada Test Range and Surrounding Areas . . . . . . . . . . . . . . . . . . . . . 44 4.1. Solid Rocket ... Rocket Mass Ratio . . . 62 4.6. Range of an ICBM from Current Missile Bases . . . . . . . . . . . . . . . . 64 4.7. Range of an ICBM from Expanded...38 4.1. Specific Impulse of Various Rocket Propellants

Project NEO Specific Impulse Testing Solutions

NASA Technical Reports Server (NTRS)

Baffa, Bill

2018-01-01

The Neo test stand is currently configured to fire a horizontally mounted rocket motor with up to 6500 lbf thrust. Currently, the Neo test stand can measure flow of liquid propellant and oxidizer, pressures residing in the closed system up to the combustion chamber. The current configuration does not have the ability to provide all data needed to compute specific impulse. This presents three methods to outfit the NEO test fixture with instrumentation allowing for calculation of specific impulse.

Theoretical Rocket Performance of Liquid Methane with Several Fluorine-Oxygen Mixtures Assuming Frozen Composition

NASA Technical Reports Server (NTRS)

Gordon, Sanford; Kastner, Michael E

1958-01-01

Theoretical rocket performance for frozen composition during expansion was calculated for liquid methane with several fluorine-oxygen mixtures for a range of pressure ratios and oxidant-fuel ratios. The parameters included are specific impulse, combustion-chamber temperature, nozzle-exit temperature molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, and thermal conductivity. The maximum calculated value of specific impulse for a chamber pressure of 600 pounds per square inch absolute (40.827atm) and an exit pressure of 1 atmosphere is 315.3 for 79.67 percent fluorine in the oxidant.

Solar thermal rocket engine (STRE) thrust characteristics at the change of engine operation mode and of the flight vehicle attitude in the solar system

NASA Astrophysics Data System (ADS)

Kudrin, O. I.

1993-10-01

Relationships are presented which describe changes in the thrust and specific impulse of a solar thermal rocket engine due to a change in the flow rate of the working fluid (hydrogen). Expressions are also presented which describe the variation of the STRE thrust and specific impulse with the distance between the flight vehicle and the sun. Results of calculations are presented for an STRE with afterburning of the working fluid (hydrogen + oxygen) using hydrogen heating by solar energy to a temperature of 2360 K.

Ignition and Performance Tests of Rocket-Based Combined Cycle Propulsion System

NASA Technical Reports Server (NTRS)

Anderson, William E.

2005-01-01

The ground testing of a Rocket Based Combined Cycle engine implementing the Simultaneous Mixing and Combustion scheme was performed at the direct-connect facility of Purdue University's High Pressure Laboratory. The fuel-rich exhaust of a JP-8/H2O2 thruster was mixed with compressed, metered air in a constant area, axisymmetric duct. The thruster was similar in design and function to that which will be used in the flight test series of Dryden's Ducted-Rocket Experiment. The determination of duct ignition limits was made based on the variation of secondary air flow rates and primary thruster equivalence ratios. Thrust augmentation and improvements in specific impulse were studied along with the pressure and temperature profiles of the duct to study mixing lengths and thermal choking. The occurrence of ignition was favored by lower rocket equivalence ratios. However, among ignition cases, better thrust and specific impulse performance were seen with higher equivalence ratios owing to the increased fuel available for combustion. Thrust and specific impulse improvements by factors of 1.2 to 1.7 were seen. The static pressure and temperature profiles allowed regions of mixing and heat addition to be identified. The mixing lengths were found to be shorter at lower rocket equivalence ratios. Total pressure measurements allowed plume-based calculation of thrust, which agreed with load-cell measured values to within 6.5-8.0%. The corresponding Mach Number profile indicated the flow was not thermally choked for the highest duct static pressure case.

Theoretical performance of liquid hydrogen and liquid fluorine as a rocket propellant

NASA Technical Reports Server (NTRS)

Gordon, Sanford; Huff, Vearl N

1953-01-01

Theoretical values of performance parameters for liquid hydrogen and liquid fluorine as a rocket propellant were calculated on the assumption of equilibrium composition during the expansion process for a wide range of fuel-oxidant and expansion ratios. The parameters included were specific impulse, combustion-chamber temperature, nozzle-exit temperature, equilibrium composition, mean molecular weight, characteristic velocity, coefficient of thrust, ration of nozzle-exit area to throat area, specific heat at constant pressure, coefficient of viscosity, and coefficient of thermal conductivity. The maximum value of specific impulse was 364.6 pound-seconds per pound for a chamber pressure of 300 pounds per square inch absolute (20.41 atm) and an exit pressure of 1 atmosphere.

A General Method for Automatic Computation of Equilibrium Compositions and Theoretical Rocket Performance of Propellants

NASA Technical Reports Server (NTRS)

Gordon, Sanford; Zeleznik, Frank J.; Huff, Vearl N.

1959-01-01

A general computer program for chemical equilibrium and rocket performance calculations was written for the IBM 650 computer with 2000 words of drum storage, 60 words of high-speed core storage, indexing registers, and floating point attachments. The program is capable of carrying out combustion and isentropic expansion calculations on a chemical system that may include as many as 10 different chemical elements, 30 reaction products, and 25 pressure ratios. In addition to the equilibrium composition, temperature, and pressure, the program calculates specific impulse, specific impulse in vacuum, characteristic velocity, thrust coefficient, area ratio, molecular weight, Mach number, specific heat, isentropic exponent, enthalpy, entropy, and several thermodynamic first derivatives.

Theoretical performance of liquid ammonia and liquid fluorine as a rocket propellant

NASA Technical Reports Server (NTRS)

Gordon, Sanford; Huff, Vearl N

1953-01-01

Theoretical values of performance parameters for liquid ammonia and liquid fluorine as a rocket propellant were calculated on the assumption of equilibrium composition during the expansion process for a wide range of fuel-oxidant and expansion ratios. The parameters included were specific impulse, combustion chamber temperature, nozzle-exit temperature, equilibrium composition, mean molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, coefficient of viscosity, and coefficient of thermal conductivity. The maximum value of specific impulse was 311.5 pound-seconds per pound for a chamber pressure of 300 pounds per square inch absolute (20.41 atm) and an exit pressure of 1 atmosphere.

Investigation of very high energy rockets for future SSTO vehicles

NASA Astrophysics Data System (ADS)

Froning, H. D., Jr.

1989-04-01

Several new propulsion possibilities are being explored in the U.S.A. which might significantly increase the amount of payload that can be propelled into orbit for a given launch vehicle weight. As such, they might enable significant reduction in the future cost of transportation between earth and space. One possibility is the combustion of matter that is in an excited atomic or molecular state. Another possibility is the annihilation of matter by means of anti-matter (by matter with identical mass and opposite electrical charge). And if an appreciable fraction of the energies released by either of these processes could be converted into the useful kinetic energy of a rocket's exhaust, a 2-6-fold increase in its specific impulse might be achieved. This paper shows that a 2-6-fold increase in rocket specific impulse might enable a 4-12-fold reduction in aerospace vehicle weight. It also shows that the specific impulse potential of excited matter or anti-matter fuels might enable transport of heavy payloads into earth orbit by means of single-stage-to-orbit vehicles that would be no heavier than current commercial airline jets.

Design of a High Temperature Radiator for the Variable Specific Impulse Magnetoplasma Rocket

NASA Technical Reports Server (NTRS)

Sheth, Rubik B.; Ungar, Eugene K.; Chambliss, Joe P.

2012-01-01

The Variable Specific Impulse Magnetoplasma Rocket (VASIMR), currently under development by Ad Astra Rocket Company (Webster, TX), is a unique propulsion system that could change the way space propulsion is performed. VASIMR's efficiency, when compared to that of a conventional chemical rocket, reduces the propellant needed for exploration missions by a factor of 10. Currently plans include flight tests of a 200 kW VASIMR system, titled VF-200, on the International Space Station (ISS). The VF-200 will consist of two 100 kW thruster units packaged together in one engine bus. Each thruster core generates 27 kW of waste heat during its 15 minute firing time. The rocket core will be maintained between 283 and 573 K by a pumped thermal control loop. The design of a high temperature radiator is a unique challenge for the vehicle design. This paper will discuss the path taken to develop a steady state and transient-based radiator design. The paper will describe the radiator design option selected for the VASIMR thermal control system for use on ISS, and how the system relates to future exploration vehicles.

Effect of ITE and nozzle exit cone erosion on specific impulse of solid rocket motors

NASA Astrophysics Data System (ADS)

Smith-Kent, Randall; Ridder, Jeffrey P.; Loh, Hai-Tien; Abel, Ralph

1993-06-01

Specific impulse loss due to the use of a slowly eroding integral throat entrance, or a throat insert, with a faster eroding nozzle exit cone is studied. It is suggested that an oblique shock wave produced by step-off erosion results in loss of specific impulse. This is studied by use of a shock capturing CFD method. The shock loss predictions for first-stage Peacekeeper and Castor 25 motors are found to match the trends of the test data. This work suggests that a loss mechanism, previously unaccounted, should be considered in the specific impulse prediction procedure for nozzles with step-off exit cone erosion.

Theoretical Performance of Liquid Hydrogen with Liquid Oxygen as a Rocket Propellant

NASA Technical Reports Server (NTRS)

Gordon, Sanford; McBride, Bonnie J.

1959-01-01

Theoretical rocket performance for both equilibrium and frozen composition during expansion was calculated for the propellant combination liquid hydrogen and liquid oxygen at four chamber pressures (60, 150, 300, and 600 lb/sq in. abs) and a wide range of pressure ratios (1 to 4000) and oxidant-fuel ratios (1.190 to 39.683). Data are given to estimate performance parameters at chamber pressures other than those for which data are tabulated. The parameters included are specific impulse, specific impulse in vacuum, combustion-chamber temperature, nozzle-exit temperature, molecular weight, molecular-weight derivatives, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, thermal conductivity, Mach number, and equilibrium gas compositions.

Analysis of a Rocket Based Combined Cycle Engine during Rocket Only Operation

NASA Technical Reports Server (NTRS)

Smith, T. D.; Steffen, C. J., Jr.; Yungster, S.; Keller, D. J.

1998-01-01

The all rocket mode of operation is a critical factor in the overall performance of a rocket based combined cycle (RBCC) vehicle. However, outside of performing experiments or a full three dimensional analysis, there are no first order parametric models to estimate performance. As a result, an axisymmetric RBCC engine was used to analytically determine specific impulse efficiency values based upon both full flow and gas generator configurations. Design of experiments methodology was used to construct a test matrix and statistical regression analysis was used to build parametric models. The main parameters investigated in this study were: rocket chamber pressure, rocket exit area ratio, percent of injected secondary flow, mixer-ejector inlet area, mixer-ejector area ratio, and mixer-ejector length-to-inject diameter ratio. A perfect gas computational fluid dynamics analysis was performed to obtain values of vacuum specific impulse. Statistical regression analysis was performed based on both full flow and gas generator engine cycles. Results were also found to be dependent upon the entire cycle assumptions. The statistical regression analysis determined that there were five significant linear effects, six interactions, and one second-order effect. Two parametric models were created to provide performance assessments of an RBCC engine in the all rocket mode of operation.

Performance of an Axisymmetric Rocket Based Combined Cycle Engine During Rocket Only Operation Using Linear Regression Analysis

NASA Technical Reports Server (NTRS)

Smith, Timothy D.; Steffen, Christopher J., Jr.; Yungster, Shaye; Keller, Dennis J.

1998-01-01

The all rocket mode of operation is shown to be a critical factor in the overall performance of a rocket based combined cycle (RBCC) vehicle. An axisymmetric RBCC engine was used to determine specific impulse efficiency values based upon both full flow and gas generator configurations. Design of experiments methodology was used to construct a test matrix and multiple linear regression analysis was used to build parametric models. The main parameters investigated in this study were: rocket chamber pressure, rocket exit area ratio, injected secondary flow, mixer-ejector inlet area, mixer-ejector area ratio, and mixer-ejector length-to-inlet diameter ratio. A perfect gas computational fluid dynamics analysis, using both the Spalart-Allmaras and k-omega turbulence models, was performed with the NPARC code to obtain values of vacuum specific impulse. Results from the multiple linear regression analysis showed that for both the full flow and gas generator configurations increasing mixer-ejector area ratio and rocket area ratio increase performance, while increasing mixer-ejector inlet area ratio and mixer-ejector length-to-diameter ratio decrease performance. Increasing injected secondary flow increased performance for the gas generator analysis, but was not statistically significant for the full flow analysis. Chamber pressure was found to be not statistically significant.

Study of Required Thrust Profile Determination of a Three Stages Small Launch Vehicle

NASA Astrophysics Data System (ADS)

Fariz, A.; Sasongko, R. A.; Poetro, R. E.

2018-04-01

The effect of solid rocket motor specifications, i.e. specific impulse and mass flow rate, and coast time on the thrust profile of three stages small launch vehicle is studied. Solid rocket motor specifications are collected from various small launch vehicle that had ever been in operation phase, and also from previous study. Comparison of orbital parameters shows that the radius of apocenter targeted can be approached using one combination of solid rocket motor specifications and appropriate coast time. However, the launch vehicle designed is failed to achieve the targeted orbit nor injecting the satellite to any orbit.

Experimental and analytical comparison of flowfields in a 110 N (25 lbf) H2/O2 rocket

NASA Technical Reports Server (NTRS)

Reed, Brian D.; Penko, Paul F.; Schneider, Steven J.; Kim, Suk C.

1991-01-01

A gaseous hydrogen/gaseous oxygen 110 N (25 lbf) rocket was examined through the RPLUS code using the full Navier-Stokes equations with finite rate chemistry. Performance tests were conducted on the rocket in an altitude test facility. Preliminary parametric analyses were performed for a range of mixture ratios and fuel film cooling pcts. It is shown that the computed values of specific impulse and characteristic exhaust velocity follow the trend of the experimental data. Specific impulse computed by the code is lower than the comparable test values by about two to three percent. The computed characteristic exhaust velocity values are lower than the comparable test values by three to four pct. Thrust coefficients computed by the code are found to be within two pct. of the measured values. It is concluded that the discrepancy between computed and experimental performance values could not be attributed to experimental uncertainty.

A Plasma Diagnostic Set for the Study of a Variable Specific Impulse Magnetoplasma Rocket

NASA Astrophysics Data System (ADS)

Squire, J. P.; Chang-Diaz, F. R.; Bengtson Bussell, R., Jr.; Jacobson, V. T.; Wootton, A. J.; Bering, E. A.; Jack, T.; Rabeau, A.

1997-11-01

The Advanced Space Propulsion Laboratory (ASPL) is developing a Variable Specific Impulse Magnetoplasma Rocket (VASIMR) using an RF heated magnetic mirror operated asymmetrically. We will describe the initial set of plasma diagnostics and data acquisition system being developed and installed on the VASIMR experiment. A U.T. Austin team is installing two fast reciprocating probes: a quadruple Langmuir and a Mach probe. These measure electron density and temperature profiles, electrostatic plasma fluctuations, and plasma flow profiles. The University of Houston is developing an array of 20 highly directional Retarding Potential Analyzers (RPA) for measuring ion energy distribution function profiles in the rocket plume, giving a measurement of total thrust. We have also developed a CAMAC based data acquisition system using LabView running on a Power Macintosh communicating through a 2 MB/s serial highway. We will present data from initial plasma operations and discuss future diagnostic development.

Optimization of operation conditions and configurations for solid-propellant ducted rocket combustors

NASA Astrophysics Data System (ADS)

Onn, Shing-Chung; Chiang, Hau-Jei; Hwang, Hang-Che; Wei, Jen-Ko; Cherng, Dao-Lien

1993-06-01

The dynamic behavior of a 2D turbulent mixing and combustion process has been studied numerically in the main combustion chamber of a solid-propellant ducted rocket (SDR). The mathematical model is based on the Favre-averaged conservation equations developed by Cherng (1990). Combustion efficiency, rather than specific impulse from earlier studies, is applied successfully to optimize the effects of two parameters by a multiple linear regression model. Specifically, the fuel-air equivalence ratio of the operating conditions and the air inlet location of configurations for the SDR combustor have been studied. For a equivalence ratio near the stoichiometric condition, the use of specific impulse or combustion efficiency will show similar trend in characterizing the reacting flow field in the combustor. For the overall fuel lean operating conditions, the change of combustion efficiency is much more sensitive to that of air inlet location than specific impulse does, suggesting combustion efficiency a better property than specific impulse in representing the condition toward flammability limits. In addition, the air inlet for maximum efficiency, in general, appears to be located at downstream of that for highest specific impulse. The optimal case for the effects of two parameters occurs at fuel lean condition, which shows a larger recirculation zone in front, deeper penetration of ram air into the combustor and much larger high temperature zone near the centerline of the combustor exit than those shown in the optimal case for overall equivalence ratio close to stoichiometric.

Specific Impulses Losses in Solid Propellant Rockets

DTIC Science & Technology

1974-12-17

binder -- polyvinyl, polyurethane, or polybutadiene) markedly increases performance. Aluminum is the most widely used metal since its energy properties...temperature is also used. -5- The specific impulse values calculated for a typical propellant with 16.4% aluminum are as follows: (p0 70 atm. p - 1 atm...Direct Measurement of Combuction Efficiency of Aluminum Analysis of the condensed phase enables the proportion of unburnt aluminum to be determined

Theoretical performance of some rocket propellants containing hydrogen, nitrogen, and oxygen

NASA Technical Reports Server (NTRS)

Miller, Riley O; Ordin, Paul M

1948-01-01

Theoretical performance data including nozzle-exit temperature, specific impulse, volume specific impulse and composition, temperature, and mean molecular weight of reaction products based on frozen equilibrium and isentropic expansion are presented for 13 propellant combinations at reaction pressure of 300 pounds per square inch absolute and expansion ratio of 20.4. On basis of maximum specific impulse alone, five fuels had the following order for any given oxidant: liquid hydrogen, hydrazine, liquid ammonia, and either hydrazine hydrate or hydroxylamine. Three oxidants with a given fuel had the following order: liquid ozone, liquid oxygen, and 100-percent hydrogen peroxide.

Nuclear propulsion - A vital technology for the exploration of Mars and the planets beyond

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.

1989-01-01

The physics and technology issues and performance potential of various direct thrust fission and fusion propulsion concepts are examined. Next to chemical propulsion the solid core fission thermal rocket (SCR) is the only other concept to be experimentally tested at the power (approx 1.5 to 5.0 GW) and thrust levels (approx 0.33 to 1.11 MN) required for manned Mars missions. With a specific impulse of approx 850 s, the SCR can perform various near-earth, cislunar and interplanetary missions with lower mass and cost requirements than its chemical counterpart. The gas core fission thermal rocket, with a specific power and impulse of approx 50 kW/kg and 5000 s offers the potential for quick courier trips to Mars (of about 80 days) or longer duration exploration cargo missions (lasting about 280 days) with starting masses of about 1000 m tons. Convenient transportation to the outer Solar System will require the development of magnetic and inertial fusion rockets (IFRs). Possessing specific powers and impulses of approx 100 kW/kg and 200-300 kilosecs, IFRs will usher in the era of the true Solar System class spaceship. Even Pluto will be accessible with roundtrip times of less than 2 years and starting masses of about 1500 m tons.

Nuclear propulsion: A vital technology for the exploration of Mars and the planets beyond

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.

1988-01-01

The physics and technology issues and performance potential of various direct thrust fission and fusion propulsion concepts are examined. Next to chemical propulsion the solid core fission thermal rocket (SCR) is the olny other concept to be experimentally tested at the power (approx 1.5 to 5.0 GW) and thrust levels (approx 0.33 to 1.11 MN) required for manned Mars missions. With a specific impulse of approx 850 s, the SCR can perform various near-Earth, cislunar and interplanetary missions with lower mass and cost requirements than its chemical counterpart. The gas core fission thermal rocket, with a specific power and impulse of approx 50 kW/kg and 5000 s offers the potential for quick courier trips to Mars (of about 80 days) or longer duration exploration cargo missions (lasting about 280 days) with starting masses of about 1000 m tons. Convenient transportation to the outer Solar System will require the development of magnetic and inertial fusion rockets (IFRs). Possessing specific powers and impulses of approx 100 kW/kg and 200-300 kilosecs, IFRs will usher in the era of the true Solar System class speceship. Even Pluto will be accessible with roundtrip times of less than 2 years and starting masses of about 1500 m tons.

Specific Impulse Definition for Ablative Laser Propulsion

NASA Technical Reports Server (NTRS)

Herren, Kenneth A.; Gregory, Don A.

2004-01-01

The term "specific impulse" is so ingrained in the field of rocket propulsion that it is unlikely that any fundamental argument would be taken seriously for its removal. It is not an ideal measure but it does give an indication of the amount of mass flow (mass loss/time), as in fuel rate, required to produce a measured thrust over some time period This investigation explores the implications of being able to accurately measure the ablation rate and how the language used to describe the specific impulse results may have to change slightly, and recasts the specific impulse as something that is not a time average. It is not currently possible to measure the ablation rate accurately in real time so it is generally just assumed that a constant amount of material will be removed for each laser pulse delivered The specific impulse dependence on the ablation rate is determined here as a correction to the classical textbook definition.

Fluidized-Solid-Fuel Injection Process

NASA Technical Reports Server (NTRS)

Taylor, William

1992-01-01

Report proposes development of rocket engines burning small grains of solid fuel entrained in gas streams. Main technical discussion in report divided into three parts: established fluidization technology; variety of rockets and rocket engines used by nations around the world; and rocket-engine equation. Discusses significance of specific impulse and ratio between initial and final masses of rocket. Concludes by stating three important reasons to proceed with new development: proposed engines safer; fluidized-solid-fuel injection process increases variety of solid-fuel formulations used; and development of fluidized-solid-fuel injection process provides base of engineering knowledge.

Design Analysis of a High Temperature Radiator for the Variable Specific Impulse Magnetoplasma Rocket (VASIMR)

NASA Technical Reports Server (NTRS)

Sheth, Rubik B.; Ungar, Eugene K.; Chambliss, Joe P.; Cassady, Leonard D.

2011-01-01

The Variable Specific Impulse Magnetoplasma Rocket (VASIMR), currently under development by Ad Astra Rocket Company, is a unique propulsion system that can potentially change the way space propulsion is performed. VASIMR's efficiency, when compared to that of a conventional chemical rocket, reduce propellant needed for exploration missions by a factor of 10. Currently plans include flight tests of a 200 kW VASIMR system, titled VF-200, on the International Space Station. The VF-200 will consist of two 100 kW thruster units packaged together in one engine bus. Each thruster unit has a unique heat rejection requirement of about 27 kW over a firing time of 15 minutes. In order to control rocket core temperatures, peak operating temperatures of about 300 C are expected within the thermal control loop. Design of a high temperature radiator is a unique challenge for the vehicle design. This paper will discuss the path taken to develop a steady state and transient based radiator design. The paper will describe radiator design options for the VASIMR thermal control system for use on ISS as well as future exploration vehicles.

Pretest uncertainty analysis for chemical rocket engine tests

NASA Technical Reports Server (NTRS)

Davidian, Kenneth J.

1987-01-01

A parametric pretest uncertainty analysis has been performed for a chemical rocket engine test at a unique 1000:1 area ratio altitude test facility. Results from the parametric study provide the error limits required in order to maintain a maximum uncertainty of 1 percent on specific impulse. Equations used in the uncertainty analysis are presented.

Bleed cycle propellant pumping in a gas-core nuclear rocket engine system

NASA Technical Reports Server (NTRS)

Kascak, A. F.; Easley, A. J.

1972-01-01

The performance of ideal and real staged primary propellant pumps and bleed-powered turbines was calculated for gas-core nuclear rocket engines over a range of operating pressures from 500 to 5000 atm. This study showed that for a required engine operating pressure of 1000 atm the pump work was about 0.8 hp/(lb/sec), the specific impulse penalty resulting from the turbine propellant bleed flow as about 10 percent; and the heat required to preheat the propellant was about 7.8 MN/(lb/sec). For a specific impulse above 2400 sec, there is an excess of energy available in the moderator due to the gamma and neutron heating that occurs there. Possible alternative pumping cycles are the Rankine or Brayton cycles.

Low Cost Nuclear Thermal Rocket Cermet Fuel Element Environment Testing

NASA Technical Reports Server (NTRS)

Bradley, D. E.; Mireles, O. R.; Hickman, R. R.

2011-01-01

Deep space missions with large payloads require high specific impulse and relatively high thrust to achieve mission goals in reasonable time frames.1,2 Conventional storable propellants produce average specific impulse. Nuclear thermal rockets capable of producing high specific impulse are proposed. Nuclear thermal rockets employ heat produced by fission reaction to heat and therefore accelerate hydrogen, which is then forced through a rocket nozzle providing thrust. Fuel element temperatures are very high (up to 3000 K), and hydrogen is highly reactive with most materials at high temperatures. Data covering the effects of high-temperature hydrogen exposure on fuel elements are limited.3 The primary concern is the mechanical failure of fuel elements that employ high-melting-point metals, ceramics, or a combination (cermet) as a structural matrix into which the nuclear fuel is distributed. The purpose of the testing is to obtain data to assess the properties of the non-nuclear support materials, as-fabricated, and determine their ability to survive and maintain thermal performance in a prototypical NTR reactor environment of exposure to hydrogen at very high temperatures. The fission process of the planned fissile material and the resulting heating performance is well known and does not therefore require that active fissile material be integrated in this testing. A small-scale test bed designed to heat fuel element samples via non-contact radio frequency heating and expose samples to hydrogen is being developed to assist in optimal material and manufacturing process selection without employing fissile material. This paper details the test bed design and results of testing conducted to date.

Program For Optimization Of Nuclear Rocket Engines

NASA Technical Reports Server (NTRS)

Plebuch, R. K.; Mcdougall, J. K.; Ridolphi, F.; Walton, James T.

1994-01-01

NOP is versatile digital-computer program devoloped for parametric analysis of beryllium-reflected, graphite-moderated nuclear rocket engines. Facilitates analysis of performance of engine with respect to such considerations as specific impulse, engine power, type of engine cycle, and engine-design constraints arising from complications of fuel loading and internal gradients of temperature. Predicts minimum weight for specified performance.

Method and apparatus to produce high specific impulse and moderate thrust from a fusion-powered rocket engine

DOE Office of Scientific and Technical Information (OSTI.GOV)

Cohen, Samuel A.; Pajer, Gary A.; Paluszek, Michael A.

A system and method for producing and controlling high thrust and desirable specific impulse from a continuous fusion reaction is disclosed. The resultant relatively small rocket engine will have lower cost to develop, test, and operate that the prior art, allowing spacecraft missions throughout the planetary system and beyond. The rocket engine method and system includes a reactor chamber and a heating system for heating a stable plasma to produce fusion reactions in the stable plasma. Magnets produce a magnetic field that confines the stable plasma. A fuel injection system and a propellant injection system are included. The propellant injectionmore » system injects cold propellant into a gas box at one end of the reactor chamber, where the propellant is ionized into a plasma. The propellant and fusion products are directed out of the reactor chamber through a magnetic nozzle and are detached from the magnetic field lines producing thrust.« less

Modeling and testing of a tube-in-tube separation mechanism of bodies in space

NASA Astrophysics Data System (ADS)

Michaels, Dan; Gany, Alon

2016-12-01

A tube-in-tube concept for separation of bodies in space was investigated theoretically and experimentally. The separation system is based on generation of high pressure gas by combustion of solid propellant and restricting the expansion of the gas only by ejecting the two bodies in opposite directions, in such a fashion that maximizes generated impulse. An interior ballistics model was developed in order to investigate the potential benefits of the separation system for a large range of space body masses and for different design parameters such as geometry and propellant. The model takes into account solid propellant combustion, heat losses, and gas phase chemical reactions. The model shows that for large bodies (above 100 kg) and typical separation velocities of 5 m/s, the proposed separation mechanism may be characterized by a specific impulse of 25,000 s, two order of magnitude larger than that of conventional solid rockets. It means that the proposed separation system requires only 1% of the propellant mass that would be needed for a conventional rocket for the same mission. Since many existing launch vehicles obtain such separation velocities by using conventional solid rocket motors (retro-rockets), the implementation of the new separation system design can reduce dramatically the mass of the separation system and increase safety. A dedicated experimental setup was built in order to demonstrate the concept and validate the model. The experimental results revealed specific impulse values of up to 27,000 s and showed good correspondence with the model.

A History of Welding on the Space Shuttle Main Engine (1975 to 2010)

NASA Technical Reports Server (NTRS)

Zimmerman, Frank R.; Russell, Carolyn K.

2010-01-01

The Space Shuttle Main Engine (SSME) is a high performance, throttleable, liquid hydrogen fueled rocket engine. High thrust and specific impulse (Isp) are achieved through a staged combustion engine cycle, combined with high combustion pressure (approx.3000psi) generated by the two-stage pump and combustion process. The SSME is continuously throttleable from 67% to 109% of design thrust level. The design criteria for this engine maximize performance and weight, resulting in a 7,800 pound rocket engine that produces over a half million pounds of thrust in vacuum with a specific impulse of 452/sec. It is the most reliable rocket engine in the world, accumulating over one million seconds of hot-fire time and achieving 100% flight success in the Space Shuttle program. A rocket engine with the unique combination of high reliability, performance, and reusability comes at the expense of manufacturing simplicity. Several innovative design features and fabrication techniques are unique to this engine. This is as true for welding as any other manufacturing process. For many of the weld joints it seemed mean cheating physics and metallurgy to meet the requirements. This paper will present a history of the welding used to produce the world s highest performance throttleable rocket engine.

Efficient solid rocket propulsion for access to space

NASA Astrophysics Data System (ADS)

Maggi, Filippo; Bandera, Alessio; Galfetti, Luciano; De Luca, Luigi T.; Jackson, Thomas L.

2010-06-01

Space launch activity is expected to grow in the next few years in order to follow the current trend of space exploitation for business purpose. Granting high specific thrust and volumetric specific impulse, and counting on decades of intense development, solid rocket propulsion is a good candidate for commercial access to space, even with common propellant formulations. Yet, some drawbacks such as low theoretical specific impulse, losses as well as safety issues, suggest more efficient propulsion systems, digging into the enhancement of consolidated techniques. Focusing the attention on delivered specific impulse, a consistent fraction of losses can be ascribed to the multiphase medium inside the nozzle which, in turn, is related to agglomeration; a reduction of agglomerate size is likely. The present paper proposes a model based on heterogeneity characterization capable of describing the agglomeration trend for a standard aluminized solid propellant formulation. Material microstructure is characterized through the use of two statistical descriptors (pair correlation function and near-contact particles) looking at the mean metal pocket size inside the bulk. Given the real formulation and density of a propellant, a packing code generates the material representative which is then statistically analyzed. Agglomerate predictions are successfully contrasted to experimental data at 5 bar for four different formulations.

High energy-density liquid rocket fuel performance

NASA Technical Reports Server (NTRS)

Rapp, Douglas C.

1990-01-01

A fuel performance database of liquid hydrocarbons and aluminum-hydrocarbon fuels was compiled using engine parametrics from the Space Transportation Engine Program as a baseline. Propellant performance parameters are introduced. General hydrocarbon fuel performance trends are discussed with respect to hydrogen-to-carbon ratio and heat of formation. Aluminum-hydrocarbon fuel performance is discussed with respect to aluminum metal loading. Hydrocarbon and aluminum-hydrocarbon fuel performance is presented with respect to fuel density, specific impulse, and propellant density specific impulse.

Propellant production from the Martian atmosphere

NASA Technical Reports Server (NTRS)

Bowles, J. V.; Tauber, M. E.; Anagnost, A. J.; Whittaker, T.

1992-01-01

Results are presented from a calculation of the specific impulses that can be generated through the combustion of cryogenic CO and O2 over a range of fuel/oxidizer ratios, chamber pressures, nozzle expansion ratios, freestream pressures representative of Mars, and the limiting conditions of equilibrium and frozen nozzle flow. For an expansion ratio of 80 and 100-atm. chamber pressure, a specific impulse of 298 sec was obtained; this is comparable to the best solid rocket propellants.

46 CFR 160.040-3 - Materials, construction, workmanship, and performance requirements.

Code of Federal Regulations, 2014 CFR

2014-10-01

... Appliance, Impulse-Projected Rocket Type (and Equipment) § 160.040-3 Materials, construction, workmanship...-projected rocket type line-throwing appliances and equipment shall be of good quality suitable for the... operation aboard vessels at sea. (c) Workmanship. Impulse-projected rocket type line-throwing appliances...

46 CFR 160.040-3 - Materials, construction, workmanship, and performance requirements.

Code of Federal Regulations, 2012 CFR

2012-10-01

... Appliance, Impulse-Projected Rocket Type (and Equipment) § 160.040-3 Materials, construction, workmanship...-projected rocket type line-throwing appliances and equipment shall be of good quality suitable for the... operation aboard vessels at sea. (c) Workmanship. Impulse-projected rocket type line-throwing appliances...

46 CFR 160.040-3 - Materials, construction, workmanship, and performance requirements.

Code of Federal Regulations, 2013 CFR

2013-10-01

... Appliance, Impulse-Projected Rocket Type (and Equipment) § 160.040-3 Materials, construction, workmanship...-projected rocket type line-throwing appliances and equipment shall be of good quality suitable for the... operation aboard vessels at sea. (c) Workmanship. Impulse-projected rocket type line-throwing appliances...

46 CFR 160.040-3 - Materials, construction, workmanship, and performance requirements.

Code of Federal Regulations, 2010 CFR

2010-10-01

... Appliance, Impulse-Projected Rocket Type (and Equipment) § 160.040-3 Materials, construction, workmanship...-projected rocket type line-throwing appliances and equipment shall be of good quality suitable for the... operation aboard vessels at sea. (c) Workmanship. Impulse-projected rocket type line-throwing appliances...

46 CFR 160.040-3 - Materials, construction, workmanship, and performance requirements.

Code of Federal Regulations, 2011 CFR

2011-10-01

... Appliance, Impulse-Projected Rocket Type (and Equipment) § 160.040-3 Materials, construction, workmanship...-projected rocket type line-throwing appliances and equipment shall be of good quality suitable for the... operation aboard vessels at sea. (c) Workmanship. Impulse-projected rocket type line-throwing appliances...

Dual-fuel, dual-mode rocket engine

NASA Technical Reports Server (NTRS)

Martin, James A. (Inventor)

1989-01-01

The invention relates to a dual fuel, dual mode rocket engine designed to improve the performance of earth-to-orbit vehicles. For any vehicle that operates from the earth's surface to earth orbit, it is advantageous to use two different fuels during its ascent. A high density impulse fuel, such as kerosene, is most efficient during the first half of the trajectory. A high specific impulse fuel, such as hydrogen, is most efficient during the second half of the trajectory. The invention allows both fuels to be used with a single rocket engine. It does so by adding a minimum number of state-of-the-art components to baseline single made rocket engines, and is therefore relatively easy to develop for near term applications. The novelty of this invention resides in the mixing of fuels before exhaust nozzle cooling. This allows all of the engine fuel to cool the exhaust nozzle, and allows the ratio of fuels used throughout the flight depend solely on performance requirements, not cooling requirements.

Effect of combustion-chamber pressure and nozzle expansion ratio on theoretical performance of several rocket propellant systems

NASA Technical Reports Server (NTRS)

Morrell, Virginia E

1956-01-01

Theoretical calculations of specific impulse to determine the separate effects of increasing the combustion-chamber pressure and the nozzle expansion ratio on the performance of the propellants, hydrogen-fluorine, hydrogen-oxygen, ammonia-fluorine and AN-F-58 fuel - white fuming nitric acid (95 percent). The results indicate that an increase in specific impulse obtainable with an increase in combustion-chamber pressure is almost entirely caused by the increased expansion ratio through the nozzle.

Radiation energy receiver for laser and solar propulsion systems

NASA Technical Reports Server (NTRS)

Rault, D. F. G.; Hertzberg, A.

1983-01-01

The concept of remotely heating a rocket propellant with a high intensity radiant energy flux is especially attractive due to its high specific impulse and large payload mass capabilities. In this paper, a radiation receiver-thruster which is especially suited to the particular thermodynamic and spectral characteristics of highly concentrated solar energy is proposed. In this receiver, radiant energy is volumetrically absorbed within a hydrogen gas seeded with alkali metal vapors. The alkali atoms and molecules absorb the radiant flux and, subsequently, transfer their internal excitation to hydrogen molecules through collisional quenching. It is shown that such a radiation receiver would outperform a blackbody cavity type receiver in both efficiency and maximum operating temperatures. A solar rocket equipped with such a receiver-thruster would deliver thrusts of several hundred newtons at a specific impulse of 1000 seconds.

Multiple-wavelength transmission measurements in rocket motor plumes

NASA Astrophysics Data System (ADS)

Kim, Hong-On

1991-09-01

Multiple-wavelength light transmission measurements were used to measure the mean particle size (d(sub 32)), index of refraction (m), and standard deviation of the small particles in the edge of the plume of a small solid propellant rocket motor. The results have shown that the multiple-wavelength light transmission measurement technique can be used to obtain these variables. The technique was shown to be more sensitive to changes in d(sub 32) and standard deviation (sigma) than to m. A GAP/AP/4.7 percent aluminum propellant burned at 25 atm produced particles with d32 = 0.150 +/- 0.006 microns, standard deviation = 1.50 +/- 0.04 and m = 1.63 +/- 0.13. The good correlation of the data indicated that only submicron particles were present in the edge of the plume. In today's budget conscious industry, the solid propellant rocket motor is an ideal propulsion system due to its low cost and simplicity. The major obstacle for solid rocket motors, however, is their limited specific impulse compared to airbreathing motors. One way to help overcome this limitation is to utilize metal fuel additives. Solid propellant rocket motors can achieve high specific impulse with metal fuel additives such as aluminum. Aluminum propellants also increase propellant densities and suppress transverse modes of combustion oscillations by damping the oscillations with the aluminum agglomerates in the combustion chamber.

An Experiment on Repetitive Pulse Operation of Microwave Rocket

DOE Office of Scientific and Technical Information (OSTI.GOV)

Oda, Yasuhisa; Shibata, Teppei; Komurasaki, Kimiya

2008-04-28

Microwave Rocket was operated with repetitive pulses. The microwave rocket model with forced breathing system was used. The pressure history in the thruster was measured and the thrust impulse was deduced. As a result, the impulse decreased at second pulse and impulses at latter pulses were constant. The dependence of the thrust performance on the partial filling rate of the thruster was compared to the thrust generation model based on the shock wave driven by microwave plasma. The experimental results showed good agreement to the predicted dependency.

A Flight Demonstration of Plasma Rocket Propulsion

NASA Technical Reports Server (NTRS)

Petro, Andrew

1999-01-01

The Advanced Space Propulsion Laboratory at the Johnson Space Center has been engaged in the development of a magneto-plasma rocket for several years. This type of rocket could be used in the future to propel interplanetary spacecraft. One advantageous feature of this rocket concept is the ability to vary its specific impulse so that it can be operated in a mode which maximizes propellant efficiency or a mode which maximizes thrust. This presentation will describe a proposed flight experiment in which a simple version of the rocket will be tested in space. In addition to the plasma rocket, the flight experiment will also demonstrate the use of a superconducting electromagnet, extensive use of heat pipes, and possibly the transfer of cryogenic propellant in space.

Analysis of quasi-hybrid solid rocket booster concepts for advanced earth-to-orbit vehicles

NASA Technical Reports Server (NTRS)

Zurawski, Robert L.; Rapp, Douglas C.

1987-01-01

A study was conducted to assess the feasibility of quasi-hybrid solid rocket boosters for advanced Earth-to-orbit vehicles. Thermochemical calculations were conducted to determine the effect of liquid hydrogen addition, solids composition change plus liquid hydrogen addition, and the addition of an aluminum/liquid hydrogen slurry on the theoretical performance of a PBAN solid propellant rocket. The space shuttle solid rocket booster was used as a reference point. All three quasi-hybrid systems theoretically offer higher specific impulse when compared with the space shuttle solid rocket boosters. However, based on operational and safety considerations, the quasi-hybrid rocket is not a practical choice for near-term Earth-to-orbit booster applications. Safety and technology issues pertinent to quasi-hybrid rocket systems are discussed.

Characterization of the non axial thrust generated by large solid propellant rocket motors in three axis stabilized ascent

NASA Technical Reports Server (NTRS)

Kosmann, W. J.; Dionne, E. R.; Klemetson, R. W.

1978-01-01

Nonaxial thrusts produced by solid rocket motors during three-axis stabilized attitude control have been determined from ascent experience on twenty three Burner II, Burner IIA and Block 5D-1 upper stage vehicles. A data base representing four different rocket motor designs (three spherical and one extended spherical) totaling twenty five three-axis stabilized firings is generated. Solid rocket motor time-varying resultant and lateral side force vector magnitudes, directions and total impulses, and roll torque couple magnitudes, directions, and total impulses are tabulated in the appendix. Population means and three sigma deviations are plotted. Existing applicable ground test side force and roll torque magnitudes and total impulses are evaluated and compared to the above experience data base. Within the spherical motor population, the selected AEDC ground test data consistently underestimated experienced motor side forces, roll torques and total impulses. Within the extended spherical motor population, the selected AEDC test data predicted experienced motor side forces, roll torques, and total impulses, with surprising accuracy considering the very small size of the test and experience populations.

A review of findings of a study of rocket based combined cycle engines applied to extensively axisymmetric single stage to orbit vehicles

NASA Technical Reports Server (NTRS)

Foster, Richard W.

1992-01-01

Extensively axisymmetric and non-axisymmetric Single Stage To Orbit (SSTO) vehicles are considered. The information is presented in viewgraph form and the following topics are presented: payload comparisons; payload as a percent of dry weight - a system hardware cost indicator; life cycle cost estimations; operations and support costs estimation; selected engine type; and rocket engine specific impulse calculation.

Direct measurement of the impulse in a magnetic thrust chamber system for laser fusion rocket

DOE Office of Scientific and Technical Information (OSTI.GOV)

Maeno, Akihiro; Yamamoto, Naoji; Nakashima, Hideki

2011-08-15

An experiment is conducted to measure an impulse for demonstrating a magnetic thrust chamber system for laser fusion rocket. The impulse is produced by the interaction between plasma and magnetic field. In the experiment, the system consists of plasma and neodymium permanent magnets. The plasma is created by a single-beam laser aiming at a polystyrene spherical target. The impulse is 1.5 to 2.2 {mu}Ns by means of a pendulum thrust stand, when the laser energy is 0.7 J. Without magnetic field, the measured impulse is found to be zero. These results indicate that the system for generating impulse is working.

First Stage of a Highly Reliable Reusable Launch System

NASA Technical Reports Server (NTRS)

Kloesel, Kurt J.; Pickrel, Jonathan B.; Sayles, Emily L.; Wright, Michael; Marriott, Darin; Holland, Leo; Kuznetsov, Stephen

2009-01-01

Electromagnetic launch assist has the potential to provide a highly reliable reusable first stage to a space access system infrastructure at a lower overall cost. This paper explores the benefits of a smaller system that adds the advantages of a high specific impulse air-breathing stage and supersonic launch speeds. The method of virtual specific impulse is introduced as a tool to emphasize the gains afforded by launch assist. Analysis shows launch assist can provide a 278-s virtual specific impulse for a first-stage solid rocket. Additional trajectory analysis demonstrates that a system composed of a launch-assisted first-stage ramjet plus a bipropellant second stage can provide a 48-percent gross lift-off weight reduction versus an all-rocket system. The combination of high-speed linear induction motors and ramjets is identified, as the enabling technologies and benchtop prototypes are investigated. The high-speed response of a standard 60 Hz linear induction motor was tested with a pulse width modulated variable frequency drive to 150 Hz using a 10-lb load, achieving 150 mph. A 300-Hz stator-compensated linear induction motor was constructed and static-tested to 1900 lbf average. A matching ramjet design was developed for use on the 300-Hz linear induction motor.

Nuclear propulsion: a vital technology for the exploration of Mars and the planets beyond

DOE Office of Scientific and Technical Information (OSTI.GOV)

Borowski, S.K.

1988-01-01

The physics and technology issues and performance potential of various direct thrust fission and fusion propulsion concepts are examined. Next to chemical propulsion the solid core fission thermal rocket (SCR) is the olny other concept to be experimentally tested at the power (approx 1.5 to 5.0 GW) and thrust levels (approx 0.33 to 1.11 MN) required for manned Mars missions. With a specific impulse of approx 850 s, the SCR can perform various near-Earth, cislunar and interplanetary missions with lower mass and cost requirements than its chemical counterpart. The gas core fission thermal rocket, with a specific power and impulsemore » of approx 50 kW/kg and 5000 s offers the potential for quick courier trips to Mars (of about 80 days) or longer duration exploration cargo missions (lasting about 280 days) with starting masses of about 1000 m tons. Convenient transportation to the outer Solar System will require the development of magnetic and inertial fusion rockets (IFRs). Possessing specific powers and impulses of approx 100 kW/kg and 200-300 kilosecs, IFRs will usher in the era of the true Solar System class speceship. Even Pluto will be accessible with roundtrip times of less than 2 years and starting masses of about 1500 m tons.« less

High Power Electric Propulsion Using The VASIMR VX-200: A Flight Technology Prototype

NASA Astrophysics Data System (ADS)

Bering, Edgar, III; Longmier, Benjamin; Glover, Tim; Chang-Diaz, Franklin; Squire, Jared; Brukardt, Michael

2008-11-01

The Variable Specific Impulse Magnetoplasma Rocket (VASIMR) is a high power magnetoplasma rocket, capable of Isp/thrust modulation at constant power. The plasma is produced by a helicon discharge. The bulk of the energy is added by ion cyclotron resonance heating (ICRH.) Axial momentum is obtained by adiabatic expansion of the plasma in a magnetic nozzle. Thrust/specific impulse ratio control in the VASIMR is primarily achieved by the partitioning of the RF power to the helicon and ICRH systems, with the proper adjustment of the propellant flow. Ion dynamics in the exhaust were studied using probes, gridded energy analyzers (RPA's), microwave interferometry and optical techniques. Results are summarize from high power ICRH experiments performed on the VX-100 using argon plasma during 2007, and on the VX-200 using argon plasma during 2008. The VX-100 has demonstrated ICRH antenna efficiency >90% and a total coupling efficiency of ˜75%. The rocket performance parameters inferred by integrating the moments of the ion energy distribution corresponds to a thrust of 2 N at an exhaust velocity of 20 km/s with the VX-100 device. The new VX-200 machine is described.

Improved hybrid rocket fuel

NASA Technical Reports Server (NTRS)

Dean, David L.

1995-01-01

McDonnell Douglas Aerospace, as part of its Independent R&D, has initiated development of a clean burning, high performance hybrid fuel for consideration as an alternative to the solid rocket thrust augmentation currently utilized by American space launch systems including Atlas, Delta, Pegasus, Space Shuttle, and Titan. It could also be used in single stage to orbit or as the only propulsion system in a new launch vehicle. Compared to solid propellants based on aluminum and ammonium perchlorate, this fuel is more environmentally benign in that it totally eliminates hydrogen chloride and aluminum oxide by products, producing only water, hydrogen, nitrogen, carbon oxides, and trace amounts of nitrogen oxides. Compared to other hybrid fuel formulations under development, this fuel is cheaper, denser, and faster burning. The specific impulse of this fuel is comparable to other hybrid fuels and is between that of solids and liquids. The fuel also requires less oxygen than similar hybrid fuels to produce maximum specific impulse, thus reducing oxygen delivery system requirements.

Reactor moderator, pressure vessel, and heat rejection system of an open-cycle gas core nuclear rocket concept

NASA Technical Reports Server (NTRS)

Taylor, M. F.; Whitmarsh, C. L., Jr.; Sirocky, P. J., Jr.; Iwanczyke, L. C.

1973-01-01

A preliminary design study of a conceptual 6000-megawatt open-cycle gas-core nuclear rocket engine system was made. The engine has a thrust of 196,600 newtons (44,200 lb) and a specific impulse of 4400 seconds. The nuclear fuel is uranium-235 and the propellant is hydrogen. Critical fuel mass was calculated for several reactor configurations. Major components of the reactor (reflector, pressure vessel, and waste heat rejection system) were considered conceptually and were sized.

Analysis and Evaluation of German Attainments and Research in the Liquid Rocket Engine Field. Volume 7. Thrust Control

DTIC Science & Technology

1951-01-01

by lowered cost, complexity, and flxed weight of the engine . An evaluation of the effect of throttling on specific impulse, as well as the way in... combustion chamber development. The throttling arrangement and the method of pump control are both closely with the design of the entire engine . As...the use of the rocket engine . For a complete coverage of these subjects, it is recommended that all volumes of this series be consulted

Design feasibility via ascent optimality for next-generation spacecraft

NASA Astrophysics Data System (ADS)

Miele, A.; Mancuso, S.

This paper deals with the optimization of the ascent trajectories for single-stage-sub-orbit (SSSO), single-stage-to-orbit (SSTO), and two-stage-to-orbit (TSTO) rocket-powered spacecraft. The maximum payload weight problem is studied for different values of the engine specific impulse and spacecraft structural factor. The main conclusions are that: feasibility of SSSO spacecraft is guaranteed for all the parameter combinations considered; feasibility of SSTO spacecraft depends strongly on the parameter combination chosen; not only feasibility of TSTO spacecraft is guaranteed for all the parameter combinations considered, but the TSTO payload is several times the SSTO payload. Improvements in engine specific impulse and spacecraft structural factor are desirable and crucial for SSTO feasibility; indeed, aerodynamic improvements do not yield significant improvements in payload. For SSSO, SSTO, and TSTO spacecraft, simple engineering approximations are developed connecting the maximum payload weight to the engine specific impulse and spacecraft structural factor. With reference to the specific impulse/structural factor domain, these engineering approximations lead to the construction of zero-payload lines separating the feasibility region (positive payload) from the unfeasibility region (negative payload).

Technology for low cost solid rocket boosters.

NASA Technical Reports Server (NTRS)

Ciepluch, C.

1971-01-01

A review of low cost large solid rocket motors developed at the Lewis Research Center is given. An estimate is made of the total cost reduction obtainable by incorporating this new technology package into the rocket motor design. The propellant, case material, insulation, nozzle ablatives, and thrust vector control are discussed. The effect of the new technology on motor cost is calculated for a typical expandable 260-in. booster application. Included in the cost analysis is the influence of motor performance variations due to specific impulse and weight changes. It is found for this application that motor costs may be reduced by up to 30% and that the economic attractiveness of future large solid rocket motors will be improved when the new technology is implemented.

Pulse Detonation Rocket Engine Research at NASA Marshall

NASA Technical Reports Server (NTRS)

Morris, Christopher I.

2003-01-01

This viewgraph representation provides an overview of research being conducted on Pulse Detonation Rocket Engines (PDRE) by the Propulsion Research Center (PRC) at the Marshall Space Flight Center. PDREs have a theoretical thermodynamic advantage over Steady-State Rocket Engines (SSREs) although unsteady blowdown processes complicate effective use of this advantage in practice; PRE is engaged in a fundamental study of PDRE gas dynamics to improve understanding of performance issues. Topics covered include: simplified PDRE cycle, comparison of PDRE and SSRE performance, numerical modeling of quasi 1-D rocket flows, time-accurate thrust calculations, finite-rate chemistry effects in nozzles, effect of F-R chemistry on specific impulse, effect of F-R chemistry on exit species mole fractions and PDRE performance optimization studies.

Nuclear thermal rocket nozzle testing and evaluation program

NASA Technical Reports Server (NTRS)

Davidian, Kenneth O.; Kacynski, Kenneth J.

1993-01-01

Performance characteristics of the Nuclear Thermal Rocket can be enhanced through the use of unconventional nozzles as part of the propulsion system. The Nuclear Thermal Rocket nozzle testing and evaluation program being conducted at the NASA Lewis is outlined and the advantages of a plug nozzle are described. A facility description, experimental designs and schematics are given. Results of pretest performance analyses show that high nozzle performance can be attained despite substantial nozzle length reduction through the use of plug nozzles as compared to a convergent-divergent nozzle. Pretest measurement uncertainty analyses indicate that specific impulse values are expected to be within + or - 1.17 pct.

Space Shuttle SRM development. [Solid Rocket Motors

NASA Technical Reports Server (NTRS)

Brinton, B. C.; Kilminster, J. C.

1979-01-01

The successful static test of the fourth Development Space Shuttle Solid Rocket Motor (SRM) in February 1979 concluded the development testing phase of the SRM Project. Qualification and flight motors are currently being fabricated, with the first qualification motor to be static tested. Delivered thrust-time traces on all development motors were very close to predicted values, and both specific and total impulse exceeded specification requirements. 'All-up' static tests conducted with a solid rocket booster equipment on development motors achieved all test objectives. Transportation and support equipment concepts have been proven, baselining is complete, and component reusability has been demonstrated. Evolution of the SRM transportation support equipment, and special test equipment designs are reviewed, and development activities discussed. Handling and processing aspects of large, heavy components are described.

Rho-Isp Revisited and Basic Stage Mass Estimating for Launch Vehicle Conceptual Sizing Studies

NASA Technical Reports Server (NTRS)

Kibbey, Timothy P.

2015-01-01

A single metric for judging between two candidate propellant combinations for a given application is sought. By using the ideal rocket equation, the essential link between propellant density and specific impulse as the two primary performance drivers can be demonstrated.

Theoretical Performance of Hydrogen-Oxygen Rocket Thrust Chambers

NASA Technical Reports Server (NTRS)

Sievers, Gilbert K.; Tomazic, William A.; Kinney, George R.

1961-01-01

Data are presented for liquid-hydrogen-liquid-oxygen thrust chambers at chamber pressures from 15 to 1200 pounds per square inch absolute, area ratios to approximately 300, and percent fuel from about 8 to 34 for both equilibrium and frozen composition during expansion. Specific impulse in vacuum, specific impulse, combustion-chamber temperature, nozzle-exit temperature, characteristic velocity, and the ratio of chamber-to-nozzle-exit pressure are included. The data are presented in convenient graphical forms to allow quick calculation of theoretical nozzle performance with over- or underexpansion, flow separation, and introduction of the propellants at various initial conditions or heat loss from the combustion chamber.

Low thrust chemical rocket technology

NASA Technical Reports Server (NTRS)

Schneider, Steven J.

1992-01-01

An on-going technology program to improve the performance of low thrust chemical rockets for spacecraft on-board propulsion applications is reviewed. Improved performance and lifetime is sought by the development of new predictive tools to understand the combustion and flow physics, introduction of high temperature materials and improved component designs to optimize performance, and use of higher performance propellants. Improved predictive technology is sought through the comparison of both local and global predictions with experimental data. Predictions are based on both the RPLUS Navier-Stokes code with finite rate kinetics and the JANNAF methodology. Data were obtained with laser-based diagnostics along with global performance measurements. Results indicate that the modeling of the injector and the combustion process needs improvement in these codes and flow visualization with a technique such as 2-D laser induced fluorescence (LIF) would aid in resolving issues of flow symmetry and shear layer combustion processes. High temperature material fabrication processes are under development and small rockets are being designed, fabricated, and tested using these new materials. Rhenium coated with iridium for oxidation protection was produced by the Chemical Vapor Deposition (CVD) process and enabled an 800 K increase in rocket operating temperature. Performance gains with this material in rockets using Earth storable propellants (nitrogen tetroxide and monomethylhydrazine or hydrazine) were obtained through component redesign to eliminate fuel film cooling and its associated combustion inefficiency while managing head end thermal soakback. Material interdiffusion and oxidation characteristics indicated that the requisite lifetimes of tens of hours were available for thruster applications. Rockets were designed, fabricated, and tested with thrusts of 22, 62, 440 and 550 N. Performance improvements of 10 to 20 seconds specific impulse were demonstrated. Higher performance propellants were evaluated: Space storable propellants, including liquid oxygen (LOX) as the oxidizer with nitrogen hydrides or hydrocarbon as fuels. Specifically, a LOX/hydrazine engine was designed, fabricated, and shown to have a 95 pct theoretical c-star which translates into a projected vacuum specific impulse of 345 seconds at an area ratio of 204:1. Further performance improvment can be obtained by the use of LOX/hydrogen propellants, especially for manned spacecraft applications, and specific designs must be developed and advanced through flight qualification.

Quasi-One-Dimensional Modeling of Pulse Detonation Rocket Engines

NASA Technical Reports Server (NTRS)

Morris, Christopher I.

2002-01-01

Pulsed detonation rocket engines (PDREs) have generated considerable research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional rocket engines. The detonative mode of combustion employed by these devices offers a thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional rocket engines and gas turbines. However, while this theoretical advantage has spurred a great deal of interest in building PDRE devices, the unsteady blowdown process intrinsic to the PDRE has made realistic estimates of the actual propulsive performance problematic. The recent review article by Kailasanath highlights some of the difficulties in comparing the available experimental measurements with numerical models. In a previous paper by the author, parametric studies of the performance of a single, straight-tube PDRE were reported. A 1-D, unsteady method of characteristics code, employing a constant-gamma assumption behind the detonation front, was developed for that study. Models of this type are computationally inexpensive, and are particularly useful for parametric performance comparisons. For example, a plot showing the specific impulse of various PDRE and steady-state rocket engine (SSRE) configurations as a function of blowdown pressure ratio. The performance curves clearly indicate that a straight-tube PDRE is superior in specific impulse to a SSRE with a sonic nozzle over the entire range of pressure ratios. Note, however, that a straight-tube PDRE in general does not compare favorably to a SSRE fitted with an optimized de Laval supersonic nozzle, particularly at the high pressure ratios typical for boost or in-space rocket applications. However, the calculations also show that if a dynamically optimized, supersonic de Laval nozzle could be could be fitted to a PDRE, then the specific impulse of the device would exceed that of a comparable SSRE. While such a nozzle is a considerable idealization, it is clear that nozzle design and optimization will play a critical role in whether the performance potential of PDREs can be effectively realized in practice. In order to study PDRE nozzle issues with greater accuracy, a quasi-one-dimensional, finite-rate chemistry CFD code has been developed by the author. Comparisons of the code with both the previous MOC model and experimental data from Stanford University are reported. The effect of constant-gamma and finite-rate chemistry assumptions on the flowfield and performance is examined. Parametric studies of the effect of nozzle throat size and expansion ratio, at various blowdown pressure ratios, are reported.

Preliminary Studies of a Pulsed Detonation Rocket Engine

NASA Technical Reports Server (NTRS)

Cambier, Jean-Luc; Adelman, H. G.; Menees, G. P.; Edwards, Thomas A. (Technical Monitor)

1995-01-01

In the new era of space exploration, there is a strong need for more efficient, cheaper and more reliable propulsion devices. With dramatic increase in specific impulse, the overall mass of fuel to be lifted into orbit is decreased, and this leads, in turn, to much lower mass requirements at lift-off, higher payload ratios and lower launch costs. The Pulsed Detonation engine (PDE) has received much attention lately due to its unique combination of simplicity, light-weight and efficiency. Current investigations focus principally on its use as a low speed, airbreathing engine, although other applications have also been proposed. Its use as a rocket propulsion device was first proposed in 1988 by the present authors. The superior efficiency of the Pulsed Detonation Rocket Engine (PDRE) is due to the near constant volume combustion process of a detonation wave. Our preliminary estimates suggest that the PDRE is theoretically capable of achieving specific impulses as high as 720 sec, a dramatic improvement over the current 480 sec of conventional rocket engines, making it competitive with nuclear thermal rockets. In addition to this remarkable efficiency, the PDRE may eliminate the need for high pressure cryogenic turbopumps, a principal source of failures. The heat transfer rates are also much lower, eliminating the need for nozzle cooling. Overall, the engine is more reliable and has a much lower weight. This paper will describe in detail the operation of the PDRE and calculate its performance, through numerical simulations. Engineering issues will be addressed and discussed, and the impact on mission profiles will also be presented. Finally, the performance of the PDRE using in-situ resources, such as CO and O2 from the martian atmosphere, will also be computed.

Experimental evaluation of a 600 lbf spacecraft rocket engine.

NASA Technical Reports Server (NTRS)

Hoehn, F. W.

1972-01-01

Experimental results are presented for a long-duration-capability (1000-sec), space-storable, bipropellant liquid rocket motor burning fluorine/hydrazine or FLOX/monomethylhydrazine. The interrelationship between injected mixture ratio and the per cent film cooling on vacuum specific impulse performance and chamber heat transfer is given. Experimental sea level measurements are used to predict space vacuum performance based upon simplified JANNAF reference procedures. Dynamic combustion stability is demonstrated over a wide range of operating conditions. Analytical results of char penetration, erosion, and ablative wall temperature distributions are presented for prototype chamber designs.

Computational design of an experimental laser-powered thruster

NASA Technical Reports Server (NTRS)

Jeng, San-Mou; Litchford, Ronald; Keefer, Dennis

1988-01-01

An extensive numerical experiment, using the developed computer code, was conducted to design an optimized laser-sustained hydrogen plasma thruster. The plasma was sustained using a 30 kW CO2 laser beam operated at 10.6 micrometers focused inside the thruster. The adopted physical model considers two-dimensional compressible Navier-Stokes equations coupled with the laser power absorption process, geometric ray tracing for the laser beam, and the thermodynamically equilibrium (LTE) assumption for the plasma thermophysical and optical properties. A pressure based Navier-Stokes solver using body-fitted coordinate was used to calculate the laser-supported rocket flow which consists of both recirculating and transonic flow regions. The computer code was used to study the behavior of laser-sustained plasmas within a pipe over a wide range of forced convection and optical arrangements before it was applied to the thruster design, and these theoretical calculations agree well with existing experimental results. Several different throat size thrusters operated at 150 and 300 kPa chamber pressure were evaluated in the numerical experiment. It is found that the thruster performance (vacuum specific impulse) is highly dependent on the operating conditions, and that an adequately designed laser-supported thruster can have a specific impulse around 1500 sec. The heat loading on the wall of the calculated thrusters were also estimated, and it is comparable to heat loading on the conventional chemical rocket. It was also found that the specific impulse of the calculated thrusters can be reduced by 200 secs due to the finite chemical reaction rate.

Space Storable Rocket Technology (SSRT) basic program

NASA Technical Reports Server (NTRS)

Chazen, M. L.; Mueller, T.; Casillas, A. R.; Huang, D.

1992-01-01

The Space Storable Rocket Technology Program (SSRT) was conducted to establish a technology for a new class of high performance and long life bipropellant engines using space storable propellants. The results are described. Task 1 evaluated several characteristics for a number of fuels to determine the best space storable fuel for use with LO2. The results indicated that LO2-N2H4 is the best propellant combination and provides the maximum mission/system capability maximum payload into GEO of satellites. Task 2 developed two models, performance and thermal. The performance model indicated the performance goal of specific impulse greater than or = 340 seconds (sigma = 204) could be achieved. The thermal model was developed and anchored to hot fire test data. Task 3 consisted of design, fabrication, and testing of a 200 lbf thrust test engine operating at a chamber pressure of 200 psia using LO2-N2H4. A total of 76 hot fire tests were conducted demonstrating performance greater than 340 (sigma = 204) which is a 25 second specific impulse improvement over the existing highest performance flight apogee type engines.

Magnetic Nozzle and Plasma Detachment Experiment

NASA Technical Reports Server (NTRS)

Chavers, Gregory; Dobson, Chris; Jones, Jonathan; Martin, Adam; Bengtson, Roger D.; Briezman, Boris; Arefiev, Alexey; Cassibry, Jason; Shuttpelz, Branwen; Deline, Christopher

2006-01-01

High power plasma propulsion can move large payloads for orbit transfer (such as the ISS), lunar missions, and beyond with large savings in fuel consumption owing to the high specific impulse. At high power, lifetime of the thruster becomes an issue. Electrodeless devices with magnetically guided plasma offer the advantage of long life since magnetic fields confine the plasma radially and keep it from impacting the material surfaces. For decades, concerns have been raised about the plasma remaining attached to the magnetic field and returning to the vehicle along the closed magnetic field lines. Recent analysis suggests that this may not be an issue of the magnetic field is properly shaped in the nozzle region and the plasma has sufficient energy density to stretch the magnetic field downstream. An experiment was performed to test the theory regarding the Magneto-hydrodynamic (MHD) detachment scenario. Data from this experiment will be presented. The Variable Specific Impulse Magnetoplasma Rocket (VASIMR) being developed by the Ad Astra Rocket Company uses a magnetic nozzle as described above. The VASIMR is also a leading candidate for exploiting an electric propulsion test platform being considered for the ISS.

Metallized Gelled Propellants: Oxygen/RP-1/aluminum Rocket Combustion Experiments

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan; Zakany, James S.

1995-01-01

A series of combustion experiments were conducted to measure the specific impulse, Cstar-, and specific-impulse efficiencies of a rocket engine using metallized gelled liquid propellants. These experiments used a small 20- to 40-1bf (89- to 178-N) thrust, modular engine consisting of an injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-wt% loadings of aluminum and gaseous oxygen was the oxidizer. Ten different injectors were used during the testing: 6 for the baseline 02/RP-1 tests and 4 for the gelled fuel tests which covered a wide range of mixture ratios. At the peak of the Isp versus oxidizer-to-fuel ratio (O/F) data, a range of 93 to 99% Cstar efficiency was reached with ungelled 02/RP-1. A Cstar efficiency range of 75 to 99% was obtained with gelled RP-l (0-wt% RP-1/Al) while the metallized 5-wt% RP-1/Al delivered a Cstar efficiency of 94 to 99% at the peak Isp in the O/F range tested. An 88 to 99% Cstar efficiency was obtained at the peak Isp of the gelled RP1/Al with 55-wt% Al. Specific impulse efficiencies for the 55-wt% RP-1/Al of 67%-83% were obtained at a 2.4:1 expansion ratio. Injector erosion was evident with the 55-wt% testing, while there was little or no erosion seen with the gelled RP-1 with 0- and 5-wt% Al. A protective layer of gelled fuel formed in the firings that minimized the damage to the rocket injector face. This effect may provide a useful technique for engine cooling. These experiments represent a first step in characterizing the performance of and operational issues with gelled RP-1 fuels.

Simplified procedures for correlation of experimentally measured and predicted thrust chamber performance

NASA Technical Reports Server (NTRS)

Powell, W. B.

1973-01-01

Thrust chamber performance is evaluated in terms of an analytical model incorporating all the loss processes that occur in a real rocket motor. The important loss processes in the real thrust chamber were identified, and a methodology and recommended procedure for predicting real thrust chamber vacuum specific impulse were developed. Simplified equations for the calculation of vacuum specific impulse are developed to relate the delivered performance (both vacuum specific impulse and characteristic velocity) to the ideal performance as degraded by the losses corresponding to a specified list of loss processes. These simplified equations enable the various performance loss components, and the corresponding efficiencies, to be quantified separately (except that interaction effects are arbitrarily assigned in the process). The loss and efficiency expressions presented can be used to evaluate experimentally measured thrust chamber performance, to direct development effort into the areas most likely to yield improvements in performance, and as a basis to predict performance of related thrust chamber configurations.

Impulse generation by detonation tubes

NASA Astrophysics Data System (ADS)

Cooper, Marcia Ann

Impulse generation with gaseous detonation requires conversion of chemical energy into mechanical energy. This conversion process is well understood in rocket engines where the high pressure combustion products expand through a nozzle generating high velocity exhaust gases. The propulsion community is now focusing on advanced concepts that utilize non-traditional forms of combustion like detonation. Such a device is called a pulse detonation engine in which laboratory tests have proven that thrust can be achieved through continuous cyclic operation. Because of poor performance of straight detonation tubes compared to conventional propulsion systems and the success of using nozzles on rocket engines, the effect of nozzles on detonation tubes is being investigated. Although previous studies of detonation tube nozzles have suggested substantial benefits, up to now there has been no systematic investigations over a range of operating conditions and nozzle configurations. As a result, no models predicting the impulse when nozzles are used exist. This lack of data has severely limited the development and evaluation of models and simulations of nozzles on pulse detonation engines. The first experimental investigation measuring impulse by gaseous detonation in plain tubes and tubes with nozzles operating in varying environment pressures is presented. Converging, diverging, and converging-diverging nozzles were tested to determine the effect of divergence angle, nozzle length, and volumetric fill fraction on impulse. The largest increases in specific impulse, 72% at an environment pressure of 100 kPa and 43% at an environment pressure of 1.4 kPa, were measured with the largest diverging nozzle tested that had a 12° half angle and was 0.6 m long. Two regimes of nozzle operation that depend on the environment pressure are responsible for these increases and were first observed from these data. To augment this experimental investigation, all data in the literature regarding partially filled detonation tubes was compiled and analyzed with models investigating concepts of energy conservation and unsteady gas dynamics. A model to predict the specific impulse was developed for partially filled tubes. The role of finite chemical kinetics in detonation products was examined through numerical simulations of the flow in nonsteady expansion waves.

Performance and heat transfer characteristics of the laser-heated rocket - A future space transportation system

NASA Technical Reports Server (NTRS)

Shoji, J. M.; Larson, V. R.

1976-01-01

The application of advanced liquid-bipropellant rocket engine analysis techniques has been utilized for prediction of the potential delivered performance and the design of thruster wall cooling schemes for laser-heated rocket thrusters. Delivered specific impulse values greater than 1000 lbf-sec/lbm are potentially achievable based on calculations for thrusters designed for 10-kW and 5000-kW laser beam power levels. A thruster wall-cooling technique utilizing a combination of regenerative cooling and a carbon-seeded hydrogen boundary layer is presented. The flowing carbon-seeded hydrogen boundary layer provides radiation absorption of the heat radiated from the high-temperature plasma. Also described is a forced convection thruster wall cooling design for an experimental test thruster.

Development and Implementation of a Scramjet Cycle Analysis Code with a Finite-Rate-Chemistry Combustion Model for Use on a Personal Computer

DTIC Science & Technology

1993-12-01

2 3 9 V List of Fi-ures Figure 1 - Functional...Block Diagram of a Scramjet ........................................ 9 Figure 2 - ’Corrected’ Specific Impulse of Hydrogen-Oxygen Rocket ............. 35...38 Figure 8 - Schematic of Northam/Anderson Mixing Model ............................ 39 Figure 9 - Pressure-Area

Assessment of infrasound signals recorded on seismic stations and infrasound arrays in the western United States using ground truth sources

NASA Astrophysics Data System (ADS)

Park, Junghyun; Hayward, Chris; Stump, Brian W.

2018-06-01

Ground truth sources in Utah during 2003-2013 are used to assess the contribution of temporal atmospheric conditions to infrasound detection and the predictive capabilities of atmospheric models. Ground truth sources consist of 28 long duration static rocket motor burn tests and 28 impulsive rocket body demolitions. Automated infrasound detections from a hybrid of regional seismometers and infrasound arrays use a combination of short-term time average/long-term time average ratios and spectral analyses. These detections are grouped into station triads using a Delaunay triangulation network and then associated to estimate phase velocity and azimuth to filter signals associated with a particular source location. The resulting range and azimuth distribution from sources to detecting stations varies seasonally and is consistent with predictions based on seasonal atmospheric models. Impulsive signals from rocket body detonations are observed at greater distances (>700 km) than the extended duration signals generated by the rocket burn test (up to 600 km). Infrasound energy attenuation associated with the two source types is quantified as a function of range and azimuth from infrasound amplitude measurements. Ray-tracing results using Ground-to-Space atmospheric specifications are compared to these observations and illustrate the degree to which the time variations in characteristics of the observations can be predicted over a multiple year time period.

High Pressure Earth Storable Rocket Technology Program-Hipes Options 1/2 Report

NASA Technical Reports Server (NTRS)

Chazen, M. L.; Sicher, D.; Calvignac, J.; Ono, D.

1999-01-01

Under the High Pressure Earth Storable Rocket Technology (HIPES) Program, TRW successfully completed testing of two 100 lbf thrust class rhenium chambers using N204-MMH. The first chamber was successfully fired for 4789 seconds of operating time with a maximum duration of 700 seconds. This chamber had been previously fired for 5230 seconds with N2O4-N2H4. The second chamber was successfully fired for 8085 seconds with a maximum firing duration of 1200 seconds. The Isp (specific impulse) for both chambers ranged from 323 lbf-sec/lbm to 330 lbf-sec/lbm.

Experimental performance of a high-area-ratio rocket nozzle at high combustion chamber pressure

NASA Technical Reports Server (NTRS)

Jankovsky, Robert S.; Kazaroff, John M.; Pavli, Albert J.

1996-01-01

An experimental investigation was conducted to determine the thrust coefficient of a high-area-ratio rocket nozzle at combustion chamber pressures of 12.4 to 16.5 MPa (1800 to 2400 psia). A nozzle with a modified Rao contour and an expansion area ratio of 1025:1 was tested with hydrogen and oxygen at altitude conditions. The same nozzle, truncated to an area ratio of 440:1, was also tested. Values of thrust coefficient are presented along with characteristic exhaust velocity efficiencies, nozzle wall temperatures, and overall thruster specific impulse.

Tailoff thrust and impulse imbalance between pairs of Space Shuttle solid rocket motors

NASA Technical Reports Server (NTRS)

Jacobs, E. P.; Yeager, J. M.

1975-01-01

The tailoff thrust and impulse imbalance between pairs of solid rocket motors is of particular interest for the Space Shuttle Vehicle because of the potential control problems that exist with this asymmetric configuration. Although a similar arrangement of solid rocket motors was utilized for the Titan Program, they produced less than one-half the thrust level of the Space Shuttle at web action time, and the overall vehicle was symmetric. Since the Titan Program does provide the most applicable actual test data, 23 flight pairs were analyzed to determine the actual tailoff thrust and impulse imbalance experienced. The results were scaled up using the predicted web action time thrust and tailoff time to arrive at values for the Space Shuttle. These values were then statistically treated to obtain a prediction of the maximum imbalance one could expect to experience during the Shuttle Program.

Grain Propellant Optimization Using Real Code Genetic Algorithm (RCGA)

NASA Astrophysics Data System (ADS)

Farizi, Muhammad Farraz Al; Oktovianus Bura, Romie; Fajar Junjunan, Soleh; Jihad, Bagus H.

2018-04-01

Grain propellant design is important in rocket motor design. The total impulse and ISP of the rocket motor is influenced by the grain propellant design. One way to get a grain propellant shape that generates the maximum total impulse value is to use the Real Code Genetic Algorithm (RCGA) method. In this paper RCGA is applied to star grain Rx-450. To find burn area of propellant used analytical method. While the combustion chamber pressures are sought with zero-dimensional equations. The optimization result can reach the desired target and increase the total impulse value by 3.3% from the initial design of Rx-450.

A Microwave Thruster for Spacecraft Propulsion

DOE Office of Scientific and Technical Information (OSTI.GOV)

Chiravalle, Vincent P

This presentation describes how a microwave thruster can be used for spacecraft propulsion. A microwave thruster is part of a larger class of electric propulsion devices that have higher specific impulse and lower thrust than conventional chemical rocket engines. Examples of electric propulsion devices are given in this presentation and it is shown how these devices have been used to accomplish two recent space missions. The microwave thruster is then described and it is explained how the thrust and specific impulse of the thruster can be measured. Calculations of the gas temperature and plasma properties in the microwave thruster aremore » discussed. In addition a potential mission for the microwave thruster involving the orbit raising of a space station is explored.« less

A Nuclear Cryogenic Propulsion Stage for Near-Term Space Missions

NASA Technical Reports Server (NTRS)

Houts, Michael G.; Kim, Tony; Emrich, William J.; Hickman, Robert R.; Broadway, Jeramie W.; Gerrish, Harold P.; Doughty, Glen E.; Adams, Robert B.; Bechtel, Ryan D.; Borowski, Stanley K.;

Recent Progress on the VASIMR Engine

NASA Technical Reports Server (NTRS)

Chang-Diaz, F. R.

2004-01-01

The development of the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) was initiated in the late 1970s to address a critical requirement for fast, high-power interplanetary space transportation. Its high-power and electrodeless design arises from the use of radio frequency (RF) waves to create and accelerate plasma in a magnetic nozzle. While not being a fusion rocket, it borrows heavily from that technology and takes advantage of the natural topology of open-ended magnetic systems. In addition the system lends itself well for Constant Power Throttling (CPT,) an important ability to vary thrust and specific impulse, over a wide operational range, while maintaining maximum power. This allows in-flight mission-optimization of thrust and specific impulse to enhance performance and reduce trip time. A NASA-led, research team, involving industry, academia and government facilities is pursuing the development of this concept in the United States. The technology can be validated, in the near term, in venues such as the International Space Station, where it can also serve as both a drag compensation device and a plasma contactor for the orbital facility. Recent advances in the development of this technology involve the demonstration of efficient propellant utilization in a flowing helicon plasma discharge as well as the experimental verification of single-pass ion acceleration, as predicted by theory I, by coupling RF power to the plasma through ion cyclotron resonance. This paper outlines these and other progress in our understanding of VASIMR physics and presents the concepts for its potential application in NASA's new vision of space exploration.

Computational Fluid Dynamics Analysis Method Developed for Rocket-Based Combined Cycle Engine Inlet

NASA Technical Reports Server (NTRS)

1997-01-01

Renewed interest in hypersonic propulsion systems has led to research programs investigating combined cycle engines that are designed to operate efficiently across the flight regime. The Rocket-Based Combined Cycle Engine is a propulsion system under development at the NASA Lewis Research Center. This engine integrates a high specific impulse, low thrust-to-weight, airbreathing engine with a low-impulse, high thrust-to-weight rocket. From takeoff to Mach 2.5, the engine operates as an air-augmented rocket. At Mach 2.5, the engine becomes a dual-mode ramjet; and beyond Mach 8, the rocket is turned back on. One Rocket-Based Combined Cycle Engine variation known as the "Strut-Jet" concept is being investigated jointly by NASA Lewis, the U.S. Air Force, Gencorp Aerojet, General Applied Science Labs (GASL), and Lockheed Martin Corporation. Work thus far has included wind tunnel experiments and computational fluid dynamics (CFD) investigations with the NPARC code. The CFD method was initiated by modeling the geometry of the Strut-Jet with the GRIDGEN structured grid generator. Grids representing a subscale inlet model and the full-scale demonstrator geometry were constructed. These grids modeled one-half of the symmetric inlet flow path, including the precompression plate, diverter, center duct, side duct, and combustor. After the grid generation, full Navier-Stokes flow simulations were conducted with the NPARC Navier-Stokes code. The Chien low-Reynolds-number k-e turbulence model was employed to simulate the high-speed turbulent flow. Finally, the CFD solutions were postprocessed with a Fortran code. This code provided wall static pressure distributions, pitot pressure distributions, mass flow rates, and internal drag. These results were compared with experimental data from a subscale inlet test for code validation; then they were used to help evaluate the demonstrator engine net thrust.

Energy Considerations of Hypothetical Space Drives

NASA Technical Reports Server (NTRS)

Millis, Marc G.

2007-01-01

The energy requirements of hypothetical, propellant-less space drives are compared to rockets. This serves to provide introductory estimates for potential benefits and to suggest analytical approaches for further study. A "space drive" is defined as an idealized form of propulsion that converts stored potential energy directly into kinetic energy using only the interactions between the spacecraft and its surrounding space. For Earth-to-orbit, the space drive uses 3.7 times less energy. For deep space travel, energy is proportional to the square of delta-v, whereas rocket energy scales exponentially. This has the effect of rendering a space drive 150-orders-of-magnitude better than a 17,000-s Specific Impulse rocket for sending a modest 5000 kg probe to traverse 5 ly in 50 years. Indefinite levitation, which is impossible for a rocket, could conceivably require 62 MJ/kg for a space drive. Assumption sensitivities and further analysis options are offered to guide further inquires.

Teaching about Impulse and Momentum

ERIC Educational Resources Information Center

Franklin, Bill

2004-01-01

This American Association of Physics Teachers/Physics Teaching Resource Agents (APPT/PTRA) spiral-bound manual features labs and demos physics teachers can use to give students hands-on opportunities to learn about impulse and momentum. "Make-and-take activities" include AAPT Apparatus Contest winners "An Air Impulse Rocket," "A Fan Driven…

A Note on Rocket Performance Comparison Through Impulse and Thrust Coefficients

NASA Astrophysics Data System (ADS)

Taylor, N. V.

Comparison of rocket motor systems is important when generating data to be used in making design decisions. In order to present meaningful comparisons, non-dimensional numbers related to performance are beneficial, as they remove effects of scale. Traditionally thrust coefficients and C* have been used to quantify the aerodynamic and chemical performance of a system respectively. However, it is argued here that in fact the thrust coefficient does not fully account for aerodynamic performance, as the impact of non-uniform flow at the throat is not accounted for. This discharge coefficient is usually allocated to the chemical efficiency through a correction to C*. However, this causes a coupling between chemical and aerodynamic efficiencies which may lead to poor design decisions. Through the use of a specific impulse coefficient, this risk is avoided, and furthermore comparison of unconventional nozzles becomes more straightforward. It is admitted, however, that this has no actual impact on real motor performance, being more in the way of a tidier `accounting' system.

Laser-heated rocket studies

NASA Technical Reports Server (NTRS)

Kemp, N. H.; Root, R. G.; Wu., P. K. S.; Caledonia, G. E.; Pirri, A. N.

1976-01-01

CW laser heated rocket propulsion was investigated in both the flowing core and stationary core configurations. The laser radiation considered was 10.6 micrometers, and the working gas was unseeded hydrogen. The areas investigated included initiation of a hydrogen plasma capable of absorbing laser radiation, the radiation emission properties of hot, ionized hydrogen, the flow of hot hydrogen while absorbing and radiating, the heat losses from the gas and the rocket performance. The stationary core configuration was investigated qualitatively and semi-quantitatively. It was found that the flowing core rockets can have specific impulses between 1,500 and 3,300 sec. They are small devices, whose heating zone is only a millimeter to a few centimeters long, and millimeters to centimeters in radius, for laser power levels varying from 10 to 5,000 kW, and pressure levels of 3 to 10 atm. Heat protection of the walls is a vital necessity, though the fraction of laser power lost to the walls can be as low as 10% for larger powers, making the rockets thermally efficient.

Airbreathing engine selection criteria for SSTO propulsion system

NASA Astrophysics Data System (ADS)

Ohkami, Yoshiaki; Maita, Masataka

1995-02-01

This paper presents airbreathing engine selection criteria to be applied to the propulsion system of a Single Stage To Orbit (SSTO). To establish the criteria, a relation among three major parameters, i.e., delta-V capability, weight penalty, and effective specific impulse of the engine subsystem, is derived as compared to these parameters of the LH2/LOX rocket engine. The effective specific impulse is a function of the engine I(sub sp) and vehicle thrust-to-drag ratio which is approximated by a function of the vehicle velocity. The weight penalty includes the engine dry weight, cooling subsystem weight. The delta-V capability is defined by the velocity region starting from the minimum operating velocity up to the maximum velocity. The vehicle feasibility is investigated in terms of the structural and propellant weights, which requires an iteration process adjusting the system parameters. The system parameters are computed by iteration based on the Newton-Raphson method. It has been concluded that performance in the higher velocity region is extremely important so that the airbreathing engines are required to operate beyond the velocity equivalent to the rocket engine exhaust velocity (approximately 4500 m/s).

The Evolution of the VASIMR Engine

NASA Technical Reports Server (NTRS)

Chang-Diaz, F. R.; Squire, Jared P.; Petro, Andrew; Nguyen, Tri X.

2001-01-01

Our future deep space explorers face many daunting challenges but three of these loom high above the rest: Physiological debilitation, radiation sickness and psychological stress. Many countermeasures are presently being considered to ameliorate these difficulties however, in the long run, two important new developments are required: abundant space power and advanced propulsion. The development of the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) addresses these important areas of need. The VASIMR is a high power, radio frequency-driven magneto plasma rocket, capable of very high exhaust velocities. In addition, its unique architecture allows in-flight mission-optimization of thrust and specific impulse to enhance performance and reduce trip time. A NASA-led, research team, involving industry, academia and government facilities is pursuing the development of this concept in the United States. The technology can be validated, in the near term, in venues such as the International Space Station, where it can also serve as both a drag compensation device and a plasma contactor for the orbital facility. Other near-Earth applications in the commercial and scientific satellite sectors are also envisioned. This presentation covers the present status of the technology, plans for its near term deployment and a vision for its future evolution.

Safety and Performance Advantages of Nitrous Oxide Fuel Blends (NOFBX) Propellants for Manned and Unmanned Spaceflight Applications

NASA Astrophysics Data System (ADS)

Taylor, R.

2012-01-01

Hydrazine, N2H4, is the current workhorse monopropellant in the spacecraft industry. Although widely used since the 1960's, hydrazine is highly toxic and its specific impulse (ISP) performance of ~230s is far lower than bipropellants and solid motors. NOFBX™ monopropellants were originally developed under NASA's Mars Advanced Technology program (2004-2007) for deep space Mars missions. This work focused on characterizing various Nitrous Oxide Fuel Blend (NOFB) monopropellants which exhibited many favorable attributes to include: (1) Mono-propulsion, (2) Isp > 320s, (3) Non-toxic constituents, (4) Non-toxic effluents, (5) Low Cost, (6) High Density Specific Impulse, (7) Non-cryogenic, (8) Wide Storable Temperature Range, (9) Deeply throttlable [between 5 - 100lbs], (10) Self Pressurizing, (11) Wide Range of materials compatibility, along with many, many other benefits. All rocket propellants carry with them a history or stigma associated with either the development or implementation of that propellant and NOFBX™ is no exception. This paper examines the benefits of NOFBX™ propellants while addressing or dispelling a number of critiques N2O based propellants acquired through the decades of rocket propellant testing.

The Vasimr Engine: Project Status and Recent Accomplishments

NASA Technical Reports Server (NTRS)

ChangDiaz, Franklin R.; Squire, Jared P.; Bering, Edgar A., III; Baitty, F. Wally; Goulding, Richard H.; Bengtson, Roger D.

2004-01-01

The development of the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) was initiated in the late 1970s to address a critical requirement for fast, high-power interplanetary space transportation. While not being a fusion rocket, it nevertheless borrows heavily from that technology and takes advantage of the natural topology of open-ended magnetic systems. In addition to its high power density and high exhaust velocity, VASIMR is capable of "constant power throttling" a feature, which allows in-flight mission-optimization of thrust and specific impulse to enhance performance and reduce trip time. A NASA-led, research team, involving industry, academia and government facilities is pursuing the development of this concept in the United States. The technology can be validated, in the near term, in venues such as the International Space Station, where it can also serve as both a drag compensation device and a plasma contactor for the orbital facility. Other near-Earth applications in the commercial and scientific satellite sectors are also envisioned. This presentation covers the evolution of the VASIMR concept to its present status, as well as recent accomplishments in our understanding of the physics. Approaches and collaborative programs addressing the major technical challenges will also be presented.

Predicting the velocity and azimuth of fragments generated by the range destruction or random failure of rocket casings and tankage

NASA Technical Reports Server (NTRS)

Eck, Marshall; Mukunda, Meera

1988-01-01

A calculational method is described which provides a powerful tool for predicting solid rocket motor (SRM) casing and liquid rocket tankage fragmentation response. The approach properly partitions the available impulse to each major system-mass component. It uses the Pisces code developed by Physics International to couple the forces generated by an Eulerian-modeled gas flow field to a Lagrangian-modeled fuel and casing system. The details of the predictive analytical modeling process and the development of normalized relations for momentum partition as a function of SRM burn time and initial geometry are discussed. Methods for applying similar modeling techniques to liquid-tankage-overpressure failures are also discussed. Good agreement between predictions and observations are obtained for five specific events.

Nuclear rocket propulsion. NASA plans and progress, FY 1991

NASA Technical Reports Server (NTRS)

Clark, John S.; Miller, Thomas J.

1991-01-01

NASA has initiated planning for a technology development project for nuclear rocket propulsion systems for space explorer initiative (SEI) human and robotic missions to the moon and Mars. An interagency project is underway that includes the Department of Energy National Laboratories for nuclear technology development. The activities of the project planning team in FY 1990 and 1991 are summarized. The progress to date is discussed, and the project plan is reviewed. Critical technology issues were identified and include: (1) nuclear fuel temperature, life, and reliability; (2) nuclear system ground test; (3) safety; (4) autonomous system operation and health monitoring; and (5) minimum mass and high specific impulse.

Metallized solid rocket propellants based on AN/AP and PSAN/AP for access to space

NASA Astrophysics Data System (ADS)

Levi, S.; Signoriello, D.; Gabardi, A.; Molinari, M.; Galfetti, L.; Deluca, L. T.; Cianfanelli, S.; Klyakin, G. F.

2009-09-01

Solid rocket propellants based on dual mixes of inorganic crystalline oxidizers (ammonium nitrate (AN) and ammonium perchlorate (AP)) with binder and a mixture of micrometric-nanometric aluminum were investigated. Ammonium nitrate is a low-cost oxidizer, producing environment friendly combustion products but with lower specific impulse compared to AP. The better performance obtained with AP and the low quantity of toxic emissions obtained by using AN have suggested an interesting compromise based on a dual mixture of the two oxidizers. To improve the thermal response of raw AN, different types of phase stabilized AN (PSAN) and AN/AP co-crystals were investigated.

Nuclear rocket propulsion: NASA plans and progress - FY 1991

NASA Technical Reports Server (NTRS)

Clark, John S.; Miller, Thomas J.

1991-01-01

NASA has initiated planning for a technology development project for nuclear rocket propulsion systems for space exploration initiative (SEI) human and robotic missions to the Moon and to Mars. An interagency project is underway that includes the Department of Energy National Laboratories for nuclear technology development. The activities of the project planning team in FY 1990 and 1991 are summarized. The progress to date is discussed, and the project plan is reviewed. Critical technology issues were identified and include: (1) nuclear fuel temperature, life, and reliability; (2) nuclear system ground test; (3) safety; (4) autonomous system operation and health monitoring; and (5) minimum mass and high specific impulse.

A mini-cavity probe reactor.

NASA Technical Reports Server (NTRS)

Hyland, R. E.

1971-01-01

The mini-cavity reactor is a rocket engine concept which combines the high specific impulse from a central gaseous fueled cavity (0.6 m diam) and NERVA type fuel elements in a driver region that is external to a moderator-reflector zone to produce a compact light weight reactor. The overall dimension including a pressure vessel that is located outside of the spherical reactor is approximately 1.21 m in diameter. Specific impulses up to 2000 sec are obtainable for 220 to 890 N of thrust with pressures less than 1000 atm. Powerplant weights including a radiator for disposing of the power in the driver region are between 4600 and 32,000 kg - less than payloads of the shuttle. This reactor could also be used as a test reactor for gas-core, MHD, breeding and materials research.

A Nuclear Cryogenic Propulsion Stage for Near-Term Space Missions

NASA Technical Reports Server (NTRS)

Houts, Michael G.; Kim, Tony; Emrich, William J.; Hickman, Robert R.; Broadway, Jeramie W.; Gerrish, Harold P.; Adams, Robert B.; Bechtel, Ryan D.; Borowski, Stanley K.; George, Jeffrey A.

2013-01-01

Development efforts in the United States have demonstrated the viability and performance potential of NTP systems. For example, Project Rover (1955 - 1973) completed 22 high power rocket reactor tests. Peak performances included operating at an average hydrogen exhaust temperature of 2550 K and a peak fuel power density of 5200 MW/m3 (Pewee test), operating at a thrust of 930 kN (Phoebus-2A test), and operating for 62.7 minutes on a single burn (NRXA6 test). Results from Project Rover indicated that an NTP system with a high thrust-toweight ratio and a specific impulse greater than 900 s would be feasible. Excellent results have also been obtained by Russia. Ternary carbide fuels developed in Russia may have the potential for providing even higher specific impulses.

Rockets -- Part II.

ERIC Educational Resources Information Center

Leitner, Alfred

1982-01-01

If two rockets are identical except that one engine burns in one-tenth the time of the other (total impulse and initial fuel mass of the two engines being the same), which rocket will rise higher? Why? The answer to this question (part 1 response in v20 n6, p410, Sep 1982) is provided. (Author/JN)

MEMS-Based Solid Propellant Rocket Array Thruster

NASA Astrophysics Data System (ADS)

Tanaka, Shuji; Hosokawa, Ryuichiro; Tokudome, Shin-Ichiro; Hori, Keiichi; Saito, Hirobumi; Watanabe, Masashi; Esashi, Masayoshi

The prototype of a solid propellant rocket array thruster for simple attitude control of a 10 kg class micro-spacecraft was completed and tested. The prototype has 10×10 φ0.8 mm solid propellant micro-rockets arrayed at a pitch of 1.2 mm on a 20×22 mm substrate. To realize such a dense array of micro-rockets, each ignition heater is powered from the backside of the thruster through an electrical feedthrough which passes along a propellant cylinder wall. Boron/potassium nitrate propellant (NAB) is used with/without lead rhodanide/potassium chlorate/nitrocellulose ignition aid (RK). Impulse thrust was measured by a pendulum method in air. Ignition required electric power of at least 3 4 W with RK and 4 6 W without RK. Measured impulse thrusts were from 2×10-5 Ns to 3×10-4 Ns after the calculation of compensation for air dumping.

Comparison of theoretical and experimental thrust performance of a 1030:1 area ratio rocket nozzle at a chamber pressure of 2413 kN/sq m (350 psia)

NASA Technical Reports Server (NTRS)

Smith, Tamara A.; Pavli, Albert J.; Kacynski, Kenneth J.

1987-01-01

The Joint Army, Navy, NASA, Air Force (JANNAF) rocket-engine performance-prediction procedure is based on the use of various reference computer programs. One of the reference programs for nozzle analysis is the Two-Dimensional Kinetics (TDK) Program. The purpose of this report is to calibrate the JANNAF procedure that has been incorporated into the December 1984 version of the TDK program for the high-area-ratio rocket-engine regime. The calibration was accomplished by modeling the performance of a 1030:1 rocket nozzle tested at NASA Lewis. A detailed description of the test conditions and TDK input parameters is given. The reuslts indicate that the computer code predicts delivered vacuum specific impulse to within 0.12 to 1.9 percent of the experimental data. Vacuum thrust coefficient predictions were within + or - 1.3 percent of experimental results. Predictions of wall static pressure were within approximately + or - 5 percent of the measured values.

Heat Exchanger Design in Combined Cycle Engines

NASA Astrophysics Data System (ADS)

Webber, H.; Feast, S.; Bond, A.

Combined cycle engines employing both pre-cooled air-breathing and rocket modes of operation are the most promising propulsion system for achieving single stage to orbit vehicles. The air-breathing phase is purely for augmentation of the mission velocity required in the rocket phase and as such must be mass effective, re-using the components of the rocket cycle, whilst achieving adequate specific impulse. This paper explains how the unique demands placed on the air-breathing cycle results in the need for sophisticated thermodynamics and the use of a series of different heat exchangers to enable precooling and high pressure ratio compression of the air for delivery to the rocket combustion chambers. These major heat exchanger roles are; extracting heat from incoming air in the precooler, topping up cycle flow temperatures to maintain constant turbine operating conditions and extracting rejected heat from the power cycle via regenerator loops for thermal capacity matching. The design solutions of these heat exchangers are discussed.

Solar rocket system concept analysis

NASA Technical Reports Server (NTRS)

Boddy, J. A.

1980-01-01

The use of solar energy to heat propellant for application to Earth orbital/planetary propulsion systems is of interest because of its performance capabilities. The achievable specific impulse values are approximately double those delivered by a chemical rocket system, and the thrust is at least an order of magnitude greater than that produced by a mercury bombardment ion propulsion thruster. The primary advantage the solar heater thruster has over a mercury ion bombardment system is that its significantly higher thrust permits a marked reduction in mission trip time. The development of the space transportation system, offers the opportunity to utilize the full performance potential of the solar rocket. The requirements for transfer from low Earth orbit (LEO) to geosynchronous equatorial orbit (GEO) was examined as the return trip, GEO to LEO, both with and without payload. Payload weights considered ranged from 2000 to 100,000 pounds. The performance of the solar rocket was compared with that provided by LO2-LH2, N2O4-MMH, and mercury ion bombardment systems.

Nuclear Thermal Propulsion: Past, Present, and a Look Ahead

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.

2014-01-01

NTR: High thrust high specific impulse (2 x LOXLH2 chemical) engine uses high power density fission reactor with enriched uranium fuel as thermal power source. Reactor heat is removed using H2 propellant which is then exhausted to produce thrust. Conventional chemical engine LH2 tanks, turbo pumps, regenerative nozzles and radiation-cooled shirt extensions used -- NTR is next evolutionary step in high performance liquid rocket engines.

Atomic hydrogen propellants: Historical perspectives and future possibilities

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

1993-01-01

Atomic hydrogen, a very high density free-radical propellant, is anticipated to generate a specific impulse of 600-1500 lb-f sec/lb-mass performance; this may facilitate the development of unique launch vehicles. A development status evaluation is presently given for atomic hydrogen investigations. It is noted that breakthroughs are required in the production, storage, and transfer of atomic hydrogen, before this fuel can become a viable rocket propellant.

Liquid Engine Design: Effect of Chamber Dimensions on Specific Impulse

NASA Technical Reports Server (NTRS)

Hoggard, Lindsay; Leahy, Joe

2009-01-01

Which assumption of combustion chemistry - frozen or equilibrium - should be used in the prediction of liquid rocket engine performance calculations? Can a correlation be developed for this? A literature search using the LaSSe tool, an online repository of old rocket data and reports, was completed. Test results of NTO/Aerozine-50 and Lox/LH2 subscale and full-scale injector and combustion chamber test results were found and studied for this task. NASA code, Chemical Equilibrium with Applications (CEA) was used to predict engine performance using both chemistry assumptions, defined here. Frozen- composition remains frozen during expansion through the nozzle. Equilibrium- instantaneous chemical equilibrium during nozzle expansion. Chamber parameters were varied to understand what dimensions drive chamber C* and Isp. Contraction Ratio is the ratio of the nozzle throat area to the area of the chamber. L is the length of the chamber. Characteristic chamber length, L*, is the length that the chamber would be if it were a straight tube and had no converging nozzle. Goal: Develop a qualitative and quantitative correlation for performance parameters - Specific Impulse (Isp) and Characteristic Velocity (C*) - as a function of one or more chamber dimensions - Contraction Ratio (CR), Chamber Length (L ) and/or Characteristic Chamber Length (L*). Determine if chamber dimensions can be correlated to frozen or equilibrium chemistry.

The SKYLON Spaceplane

NASA Astrophysics Data System (ADS)

Varvill, R.; Bond, A.

SKYLON is a single stage to orbit (SSTO) winged spaceplane designed to give routine low cost access to space. At a gross takeoff weight of 275 tonnes of which 220 tonnes is propellant the vehicle is capable of placing 12 tonnes into an equatorial low Earth orbit. The vehicle configuration consists of a slender fuselage containing the propellant tankage and payload bay with delta wings located midway along the fuselage carrying the SABRE engines in axisymmetric nacelles on the wingtips. The vehicle takes off and lands horizontally on it's own undercarriage. The fuselage is constructed as a multilayer structure consisting of aeroshell, insulation, structure and tankage. SKYLON employs extant or near term materials technology in order to minimise development cost and risk. The SABRE engines have a dual mode capability. In rocket mode the engine operates as a closed cycle liquid oxygen/liquid hydrogen high specific impulse rocket engine. In airbreathing mode (from takeoff to Mach 5) the liquid oxygen flow is replaced by atmospheric air, increasing the installed specific impulse 3-6 fold. The airflow is drawn into the engine via a 2 shock axisymmetric intake and cooled to cryogenic temperatures prior to compression. The hydrogen fuel flow acts as a heat sink for the closed cycle helium loop before entering the main combustion chamber.

Outbrief - Long Life Rocket Engine Panel

NASA Technical Reports Server (NTRS)

Quinn, Jason Eugene

2004-01-01

This white paper is an overview of the JANNAF Long Life Rocket Engine (LLRE) Panel results from the last several years of activity. The LLRE Panel has met over the last several years in order to develop an approach for the development of long life rocket engines. Membership for this panel was drawn from a diverse set of the groups currently working on rocket engines (Le. government labs, both large and small companies and university members). The LLRE Panel was formed in order to determine the best way to enable the design of rocket engine systems that have life capability greater than 500 cycles while meeting or exceeding current performance levels (Specific Impulse and Thrust/Weight) with a 1/1,OOO,OOO likelihood of vehicle loss due to rocket system failure. After several meetings and much independent work the panel reached a consensus opinion that the primary issues preventing LLRE are a lack of: physics based life prediction, combined loads prediction, understanding of material microphysics, cost effective system level testing. and the inclusion of fabrication process effects into physics based models. With the expected level of funding devoted to LLRE development, the panel recommended that fundamental research efforts focused on these five areas be emphasized.

Mariner Venus/Mercury 1973 rocket engine assembly

NASA Technical Reports Server (NTRS)

Snoke, D. R.; Williams, R. S.

1972-01-01

The fabrication and test of rocket engine assemblies (REA) for Mariner Venus/Mercury 1973 are reported. The fabrication, assembly and flight acceptance test of seven REA's including the type approval test of one engine and fabrication of one additional kit consisting of detail parts for an engine ready for catalyst loading are presented. The MV/M '73 REA which is a nominal 51 lbs thrust monopropellant engine is described. Under steady state operation the specific impulse is not less than 228 lb-sec at 55 lb and 218.5 lb-sec at 10 lb thrust varying linearly between these limits. The characteristic velocity is not less than 4100 ft/sec at any thrust level.

MHD Energy Bypass Scramjet Performance with Real Gas Effects

NASA Technical Reports Server (NTRS)

Park, Chul; Mehta, Unmeel B.; Bogdanoff, David W.

2000-01-01

The theoretical performance of a scramjet propulsion system incorporating an magneto-hydro-dynamic (MHD) energy bypass scheme is calculated. The one-dimensional analysis developed earlier, in which the theoretical performance is calculated neglecting skin friction and using a sudden-freezing approximation for the nozzle flow, is modified to incorporate the method of Van Driest for turbulent skin friction and a finite-rate chemistry calculation in the nozzle. Unlike in the earlier design, in which four ramp compressions occurred in the pitch plane, in the present design the first two ramp compressions occur in the pitch plane and the next two compressions occur in the yaw plane. The results for the simplified design of a spaceliner show that (1) the present design produces higher specific impulses than the earlier design, (2) skin friction substantially reduces thrust and specific impulse, and (3) the specific impulse of the MHD-bypass system is still better than the non-MHD system and typical rocket over a narrow region of flight speeds and design parameters. Results suggest that the energy management with MHD principles offers the possibility of improving the performance of the scramjet. The technical issues needing further studies are identified.

Feasibility and Performance of the Microwave Thermal Rocket Launcher

NASA Astrophysics Data System (ADS)

Parkin, Kevin L. G.; Culick, Fred E. C.

2004-03-01

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%.

Laser rocket system analysis

NASA Technical Reports Server (NTRS)

Jones, W. S.; Forsyth, J. B.; Skratt, J. P.

1979-01-01

The laser rocket systems investigated in this study were for orbital transportation using space-based, ground-based and airborne laser transmitters. The propulsion unit of these systems utilizes a continuous wave (CW) laser beam focused into a thrust chamber which initiates a plasma in the hydrogen propellant, thus heating the propellant and providing thrust through a suitably designed nozzle and expansion skirt. The specific impulse is limited only by the ability to adequately cool the thruster and the amount of laser energy entering the engine. The results of the study showed that, with advanced technology, laser rocket systems with either a space- or ground-based laser transmitter could reduce the national budget allocated to space transportation by 10 to 345 billion dollars over a 10-year life cycle when compared to advanced chemical propulsion systems (LO2-LH2) of equal capability. The variation in savings depends upon the projected mission model.

Design and Testing of a Liquid Nitrous Oxide and Ethanol Fueled Rocket Engine

DOE Office of Scientific and Technical Information (OSTI.GOV)

Youngblood, Stewart

A small-scale, bi-propellant, liquid fueled rocket engine and supporting test infrastructure were designed and constructed at the Energetic Materials Research and Testing Center (EMRTC). This facility was used to evaluate liquid nitrous oxide and ethanol as potential rocket propellants. Thrust and pressure measurements along with high-speed digital imaging of the rocket exhaust plume were made. This experimental data was used for validation of a computational model developed of the rocket engine tested. The developed computational model was utilized to analyze rocket engine performance across a range of operating pressures, fuel-oxidizer mixture ratios, and outlet nozzle configurations. A comparative study ofmore » the modeling of a liquid rocket engine was performed using NASA CEA and Cantera, an opensource equilibrium code capable of being interfaced with MATLAB. One goal of this modeling was to demonstrate the ability of Cantera to accurately model the basic chemical equilibrium, thermodynamics, and transport properties for varied fuel and oxidizer operating conditions. Once validated for basic equilibrium, an expanded MATLAB code, referencing Cantera, was advanced beyond CEAs capabilities to predict rocket engine performance as a function of supplied propellant flow rate and rocket engine nozzle dimensions. Cantera was found to comparable favorably to CEA for making equilibrium calculations, supporting its use as an alternative to CEA. The developed rocket engine performs as predicted, demonstrating the developedMATLAB rocket engine model was successful in predicting real world rocket engine performance. Finally, nitrous oxide and ethanol were shown to perform well as rocket propellants, with specific impulses experimentally recorded in the range of 250 to 260 seconds.« less

Asymmetric Shock Wave Generation in a Microwave Rocket Using a Magnetic Field

NASA Astrophysics Data System (ADS)

Takahashi, Masayuki

2017-10-01

A plasma pattern is reproduced by coupling simulations between a particle-in- cell with Monte Carlo collisions model and a finite-difference time-domain simulation for an electromagnetic wave propagation when an external magnetic field is applied to the breakdown volume inside a microwave-rocket nozzle. The propagation speed and energy-absorption rate of the plasma are estimated based on the breakdown simulation, and these are utilized to reproduce shock wave propagation, which provides impulsive thrust for the microwave rocket. The shock wave propagation is numerically reproduced by solving the compressible Euler equation with an energy source of the microwave heating. The shock wave is asymmetrically generated inside the nozzle when the electron cyclotron resonance region has a lateral offset, which generates lateral and angular impulses for postural control of the vehicle. It is possible to develop an integrated device to maintain beaming ight of the microwave rocket, achieving both axial thrust improvement and postural control, by controlling the spatial distribution of the external magnetic field.

Rapid Mars transits with exhaust-modulated plasma propulsion

NASA Technical Reports Server (NTRS)

Chang-Diaz, Franklin R.; Braden, Ellen; Johnson, Ivan; Hsu, Michael M.; Yang, Tien Fang

1995-01-01

The operational characteristics of the Exhaust-Modulated Plasma Rocket are described. Four basic human and robotic mission scenarios to Mars are analyzed using numerical optimization techniques at variable specific impulse and constant power. The device is well suited for 'split-sprint' missions, allowing fast, one-way low-payload human transits of 90 to 104 days, as well as slower, 180-day, high-payload robotic precursor flights. Abort capabilities, essential for human missions, are also explored.

Physics and potentials of fissioning plasmas for space power and propulsion

NASA Technical Reports Server (NTRS)

Thom, K.; Schwenk, F. C.; Schneider, R. T.

1976-01-01

Fissioning uranium plasmas are the nuclear fuel in conceptual high-temperature gaseous-core reactors for advanced rocket propulsion in space. A gaseous-core nuclear rocket would be a thermal reactor in which an enriched uranium plasma at about 10,000 K is confined in a reflector-moderator cavity where it is nuclear critical and transfers its fission power to a confining propellant flow for the production of thrust at a specific impulse up to 5000 sec. With a thrust-to-engine weight ratio approaching unity, the gaseous-core nuclear rocket could provide for propulsion capabilities needed for manned missions to the nearby planets and for economical cislunar ferry services. Fueled with enriched uranium hexafluoride and operated at temperatures lower than needed for propulsion, the gaseous-core reactor scheme also offers significant benefits in applications for space and terrestrial power. They include high-efficiency power generation at low specific mass, the burnup of certain fission products and actinides, the breeding of U-233 from thorium with short doubling times, and improved convenience of fuel handling and processing in the gaseous phase.

Boiler and Pressure Balls Monopropellant Thermal Rocket Engine

NASA Technical Reports Server (NTRS)

Greene, William D. (Inventor)

2009-01-01

The proposed technology is a rocket engine cycle utilizing as the propulsive fluid a low molecular weight, cryogenic fluid, typically liquid hydrogen, pressure driven, heated, and expelled through a nozzle to generate high velocity and high specific impulse discharge gas. The proposed technology feeds the propellant through the engine cycle without the use of a separate pressurization fluid and without the use of turbomachinery. Advantages of the proposed technology are found in those elements of state-of-the-art systems that it avoids. It does not require a separate pressurization fluid or a thick-walled primary propellant tank as is typically required for a classical pressure-fed system. Further, it does not require the acceptance of intrinsic reliability risks associated with the use of turbomachinery

Carbon monoxide and oxygen combustion experiments: A demonstration of Mars in situ propellants

NASA Technical Reports Server (NTRS)

Linne, Diane L.

1991-01-01

The feasibility of using carbon monoxide and oxygen as rocket propellants was examined both experimentally and theoretically. The steady-state combustion of carbon monoxide and oxygen was demonstrated for the first time in a subscale rocket engine. Measurements of experimental characteristic velocity, vacuum specific impulse, and thrust coefficient efficiency were obtained over a mixture ratio range of 0.30 to 2.0 and a chamber pressures of 1070 and 530 kPa. The theoretical performance of the propellant combination was studied parametrically over the same mixture ratio range. In addition to one dimensional ideal performance predictions, various performance reduction mechanisms were also modeled, including finite-rate kinetic reactions, two-dimensional divergence effects and viscous boundary layer effects.

Optimal trajectories for an aerospace plane. Part 1: Formulation, results, and analysis

NASA Technical Reports Server (NTRS)

Miele, Angelo; Lee, W. Y.; Wu, G. D.

1990-01-01

The optimization of the trajectories of an aerospace plane is discussed. This is a hypervelocity vehicle capable of achieving orbital speed, while taking off horizontally. The vehicle is propelled by four types of engines: turbojet engines for flight at subsonic speeds/low supersonic speeds; ramjet engines for flight at moderate supersonic speeds/low hypersonic speeds; scramjet engines for flight at hypersonic speeds; and rocket engines for flight at near-orbital speeds. A single-stage-to-orbit (SSTO) configuration is considered, and the transition from low supersonic speeds to orbital speeds is studied under the following assumptions: the turbojet portion of the trajectory has been completed; the aerospace plane is controlled via the angle of attack and the power setting; the aerodynamic model is the generic hypersonic aerodynamics model example (GHAME). Concerning the engine model, three options are considered: (EM1), a ramjet/scramjet combination in which the scramjet specific impulse tends to a nearly-constant value at large Mach numbers; (EM2), a ramjet/scramjet combination in which the scramjet specific impulse decreases monotonically at large Mach numbers; and (EM3), a ramjet/scramjet/rocket combination in which, owing to stagnation temperature limitations, the scramjet operates only at M approx. less than 15; at higher Mach numbers, the scramjet is shut off and the aerospace plane is driven only by the rocket engines. Under the above assumptions, four optimization problems are solved using the sequential gradient-restoration algorithm for optimal control problems: (P1) minimization of the weight of fuel consumed; (P2) minimization of the peak dynamic pressure; (P3) minimization of the peak heating rate; and (P4) minimization of the peak tangential acceleration.

Evaluation of an Ejector Ramjet Based Propulsion System for Air-Breathing Hypersonic Flight

NASA Technical Reports Server (NTRS)

Thomas, Scott R.; Perkins, H. Douglas; Trefny, Charles J.

1997-01-01

A Rocket Based Combined Cycle (RBCC) engine system is designed to combine the high thrust to weight ratio of a rocket along with the high specific impulse of a ramjet in a single, integrated propulsion system. This integrated, combined cycle propulsion system is designed to provide higher vehicle performance than that achievable with a separate rocket and ramjet. The RBCC engine system studied in the current program is the Aerojet strutjet engine concept, which is being developed jointly by a government-industry team as part of the Air Force HyTech program pre-PRDA activity. The strutjet is an ejector-ramjet engine in which small rocket chambers are embedded into the trailing edges of the inlet compression struts. The engine operates as an ejector-ramjet from takeoff to slightly above Mach 3. Above Mach 3 the engine operates as a ramjet and transitions to a scramjet at high Mach numbers. For space launch applications the rockets would be re-ignited at a Mach number or altitude beyond which air-breathing propulsion alone becomes impractical. The focus of the present study is to develop and demonstrate a strutjet flowpath using hydrocarbon fuel at up to Mach 7 conditions.

Combustion of metal agglomerates in a solid rocket core flow

NASA Astrophysics Data System (ADS)

Maggi, Filippo; Dossi, Stefano; DeLuca, Luigi T.

2013-12-01

The need for access to space may require the use of solid propellants. High thrust and density are appealing features for different applications, spanning from boosting phase to other service applications (separation, de-orbiting, orbit insertion). Aluminum is widely used as a fuel in composite solid rocket motors because metal oxidation increases enthalpy release in combustion chamber and grants higher specific impulse. Combustion process of metal particles is complex and involves aggregation, agglomeration and evolution of reacting particulate inside the core flow of the rocket. It is always stated that residence time should be enough in order to grant complete metal oxidation but agglomerate initial size, rocket grain geometry, burning rate, and other factors have to be reconsidered. New space missions may not require large rocket systems and metal combustion efficiency becomes potentially a key issue to understand whether solid propulsion embodies a viable solution or liquid/hybrid systems are better. A simple model for metal combustion is set up in this paper. Metal particles are represented as single drops trailed by the core flow and reacted according to Beckstead's model. The fluid dynamics is inviscid, incompressible, 1D. The paper presents parametric computations on ideal single-size particles as well as on experimental agglomerate populations as a function of operating rocket conditions and geometries.

A study of variable thrust, variable specific impulse trajectories for solar system exploration

NASA Astrophysics Data System (ADS)

Sakai, Tadashi

A study has been performed to determine the advantages and disadvantages of variable thrust and variable Isp (specific impulse) trajectories for solar system exploration. There have been several numerical research efforts for variable thrust, variable Isp, power-limited trajectory optimization problems. All of these results conclude that variable thrust, variable Isp (variable specific impulse, or VSI) engines are superior to constant thrust, constant Isp (constant specific impulse; or CSI) engines. However, most of these research efforts assume a mission from Earth to Mars, and some of them further assume that these planets are circular and coplanar. Hence they still lack the generality. This research has been conducted to answer the following questions: (1) Is a VSI engine always better than a CSI engine or a high thrust engine for any mission to any planet with any time of flight considering lower propellant mass as the sole criterion? (2) If a planetary swing-by is used for a VSI trajectory, is the fuel savings of a VSI swing-by trajectory better than that of a CSI swing-by or high thrust swing-by trajectory? To support this research, an unique, new computer-based interplanetary trajectory calculation program has been created. This program utilizes a calculus of variations algorithm to perform overall optimization of thrust, Isp, and thrust vector direction along a trajectory that minimizes fuel consumption for interplanetary travel. It is assumed that the propulsion system is power-limited, and thus the compromise between thrust and Isp is a variable to be optimized along the flight path. This program is capable of optimizing not only variable thrust trajectories but also constant thrust trajectories in 3-D space using a planetary ephemeris database. It is also capable of conducting planetary swing-bys. Using this program, various Earth-originating trajectories have been investigated and the optimized results have been compared to traditional CSI and high thrust trajectory solutions. Results show that VSI rocket engines reduce fuel requirements for any mission compared to CSI rocket engines. Fuel can be saved by applying swing-by maneuvers for VSI engines; but the effects of swing-bys due to VSI engines are smaller than that of CSI or high thrust engines.

A comprehensive cold gas performance study of the Pocket Rocket radiofrequency electrothermal microthruster

NASA Astrophysics Data System (ADS)

Ho, Teck Seng; Charles, Christine; Boswell, Roderick W.

2016-12-01

This paper presents computational fluid dynamics simulations of the cold gas operation of Pocket Rocket and Mini Pocket Rocket radiofrequency electrothermal microthrusters, replicating experiments performed in both sub-Torr and vacuum environments. This work takes advantage of flow velocity choking to circumvent the invalidity of modelling vacuum regions within a CFD simulation, while still preserving the accuracy of the desired results in the internal regions of the microthrusters. Simulated results of the plenum stagnation pressure is in precise agreement with experimental measurements when slip boundary conditions with the correct tangential momentum accommodation coefficients for each gas are used. Thrust and specific impulse is calculated by integrating the flow profiles at the exit of the microthrusters, and are in good agreement with experimental pendulum thrust balance measurements and theoretical expectations. For low thrust conditions where experimental instruments are not sufficiently sensitive, these cold gas simulations provide additional data points against which experimental results can be verified and extrapolated. The cold gas simulations presented in this paper will be used as a benchmark to compare with future plasma simulations of the Pocket Rocket microthruster.

Electric propulsion technology

NASA Technical Reports Server (NTRS)

Finke, R. C.

1980-01-01

The advanced electric propulsion program is directed towards lowering the specific impulse and increasing the thrust per unit of ion thruster systems. In addition, electrothermal and electromagnetic propulsion technologies are being developed to attempt to fill the gap between the conventional ion thruster and chemical rocket systems. Most of these new concepts are exagenous and are represented by rail accelerators, ablative Teflon thrusters, MPD arcs, Free Radicals, etc. Endogenous systems such as metallic hydrogen offer great promise and are also being pursued.

Theoretical and experimental study of a thruster discharging a weight

NASA Astrophysics Data System (ADS)

Michaels, Dan; Gany, Alon

2014-06-01

An innovative concept for a rocket type thruster that can be beneficial for spacecraft trajectory corrections and station keeping was investigated both experimentally and theoretically. It may also be useful for divert and attitude control systems (DACS). The thruster is based on a combustion chamber discharging a weight through an exhaust tube. Calculations with granular double-base propellant and a solid ejected weight reveal that a specific impulse based on the propellant mass of well above 400 s can be obtained. An experimental thruster was built in order to demonstrate the new idea and validate the model. The thruster impulse was measured both directly with a load cell and indirectly by using a pressure transducer and high speed photography of the weight as it exits the tube, with both ways producing very similar total impulse measurement. The good correspondence between the computations and the measured data validates the model as a useful tool for studying and designing such a thruster.

The Story of the Nuclear Rocket: Back to the Future

NASA Astrophysics Data System (ADS)

Dewar, James A.

2002-01-01

The United States had a nuclear rocket development program from 1955-1973 called Project Rover/NERVA. Twenty reactor tests demonstrated conclusively the superiority, flexibility and reliability of nuclear rocket engines over their chemical counterparts. This paper surveys the technical accomplishments from that perspective, to help illustrate why many call for the program's reestablishment. Most focus on the large NERVA, but this review will consider the little known Small Nuclear Engine. KIWI-B1B was one of the first tests in which nuclear rockets demonstrated their superiority. It ejected its core as it rose to 1000MW (a megawatt equals 50 pounds of thrust). This seems contradictory, how can a `failure' demonstrate superiority? Precisely in this: the reactor remained controllable going to and from 1000MW, still ejecting its core, but still turning out power. That gave insurance to a mission. A solid or liquid chemical engine suffering similar damage would likely shutdown or blow up. KIWI-TNT and Phoebus-1A had planned and unplanned accidents. That verified the safety of nuclear engines in launch operations. NRX/EST and XE-Prime proved they could startup reliably under their own power in a simulated space environment and change power without loss of specific impulse or control, from 20MW to 1000MW and back. That gave flexibility for mid-course corrections, maneuvering between orbits or breaking into orbit. Pewee and the Nuclear Furnace tested fuels to achieve 10 hours of engine operation with 60 recycles (stops and starts). That meant an engine could perform multiple missions. Work started on fuels promising1000 seconds of specific impulse. That meant increased power and payload capacity and speed. This contrasts with the 450 seconds of LOX/LH2. The NERVA of 1971 would be 1500MW, with 10/60 capability and 825 seconds of a specific impulse. Later generation NERVAs would be in excess of 1000 seconds, 3000MW and 10/60. The Nixon Administration cancelled it in 1971. After its demise, the Small Nuclear Engine appeared for unmanned missions. To fit in the space shuttle's 15 by 60 foot cargo bay, the 10 foot long engine would be 400MW, weigh 5600 pounds and use slush hydrogen. That left 50 feet and almost 60,000 pounds for the tank, propellant and payload that could vary in size, but it was nominally 5 tons. It would cost 500 million (in1972 dollars) and take a decade to develop. It had NERVA's operating characteristics, but subsequent generation systems envisioned longer engine life and recycle capability and specific impulses of 1000+ seconds. Nixon ended this in 1973. By reconsidering it instead of a nuclear electric engine that serves only space science, the nation could gain a fast, powerful system that would radically change most future unmanned space missions. With its recycle capability, a single engine could ferry large scientific payloads swiftly throughout the solar system. Yet it also could propel heavy national security and commercial payloads to geo-synchronous orbit. NASA might even offer a satellite retrieval service. Thus, one lesson is clear: it is 1960s era technology, but the Small Engine is not obsolete. If developed, it would serve not just one, but three users yet have growth potential for decades for an ever more expansive space program.

RSRM-3 (360L003) Ballistics/Mass Properties Report

NASA Technical Reports Server (NTRS)

Laubacher, B. A.; Richards, M. C.

1989-01-01

The propulsion performance and reconstructed mass properties data from Morton Thiokol's RSRM-3 motors which were assigned to the STS-29 launch are presented. The composite type solid propellant burn rates were close to predicted. The performance of the pair of motors were compared to some CEI Specifications. The performance from each motor as well as matched pair performance values were well within the CEI specification requirements. The nominal thrust time curve and impulse gate information is included. Post flight reconstructed Redesigned Solid Rocket Motor (RSRM) mass properties are within expected values for the lightweight configuration.

Basic Studies Relative to the Syntheses of Unsymmetrical Dimethylhydrazine and Monomethylhydrazine by Chloramination

DTIC Science & Technology

1979-12-07

8217 changed to faint yellow. After the chloramination, the reaction mixture was allowed to stand at room temperature to allow volatile gases to ’escape. The...energy rocketI and space shuttle fuels. These compounds have large heats of combustion , relatively :,igh densities and high specific impulses. Until...a solution of KOH in n-butanol was carried out in the presence of NH3 at room temperature . The major products were 1,1-dimethylhydrazine, the

Conceptual Launch Vehicles Using Metallic Hydrogen Propellant

NASA Astrophysics Data System (ADS)

Cole, John W.; Silvera, Isaac F.; Foote, John P.

2008-01-01

Solid molecular hydrogen is predicted to transform into an atomic solid with metallic properties under pressures >4.5 Mbar. Atomic metallic hydrogen is predicted to be metastable, limited by some critical temperature and pressure, and to store very large amounts of energy. Experiments may soon determine the critical temperature, critical pressure, and specific energy availability. It is useful to consider the feasibility of using metastable atomic hydrogen as a rocket propellant. If one assumes that metallic hydrogen is stable at usable temperatures and pressures, and that it can be affordably produced, handled, and stored, then it may be a useful rocket propellant. Assuming further that the available specific energy can be determined from the recombination of the atoms into molecules (216 MJ/kg), then conceptual engines and launch vehicle concepts can be developed. Under these assumptions, metallic hydrogen would be a revolutionary new rocket fuel with a theoretical specific impulse of 1700 s at a chamber pressure of 100 atm. A practical problem that arises is that rocket chamber temperatures may be too high for the use of this pure fuel. This paper examines an engine concept that uses liquid hydrogen or water as a diluent coolant for the metallic hydrogen to reduce the chamber temperature to usable values. Several launch vehicles are then conceptually developed. Results indicate that if metallic hydrogen is experimentally found to have the properties assumed in this analysis, then there are significant benefits. These benefits become more attractive as the chamber temperatures increase.

Advanced small rocket chambers: Option 1, 14 lbf Ir-Re rocket

NASA Technical Reports Server (NTRS)

Jassowski, Donald M.; Gage, Mark L.

1992-01-01

A high performance Ir-Re 14 lbf (62 N) chamber and nozzle which can be a direct replacement for a production engine was designed, built, hot fired and vibration acceptance tested. It passed all acceptance tests satisfactorily and demonstrated a 20 sec increase in specific impulse (Is) over the conventional 14 lbf silicide coated Cb chamber. The high performance engine uses the production valve and injector without modification. Incorporation of a secondary mixing device or Boundary Layer Trip within the combustion chamber results in elimination of the fuel film coolant, improvement in flow uniformity, the 20 sec performance increase, and reduction of a potential source of spacecraft contamination. Measured Is was 305 sec at 75:1 area ratio, with monomenthylhydrazine and nitrogen tetroxide propellants. Qualification tests remain to be done.

Solid rocket booster performance evaluation model. Volume 3: Sample case. [propellant combustion simulation/internal ballistics

NASA Technical Reports Server (NTRS)

1974-01-01

The solid rocket booster performance evaluation model (SRB-11) is used to predict internal ballistics in a sample motor. This motor contains a five segmented grain. The first segment has a 14 pointed star configuration with a web which wraps partially around the forward dome. The other segments are circular in cross-section and are tapered along the interior burning surface. Two of the segments are inhibited on the forward face. The nozzle is not assumed to be submerged. The performance prediction is broken into two simulation parts: the delivered end item specific impulse and the propellant properties which are required as inputs for the internal ballistics module are determined; and the internal ballistics for the entire burn duration of the motor are simulated.

Rho-Isp Revisited and Basic Stage Mass Estimating for Launch Vehicle Conceptual Sizing Studies

NASA Technical Reports Server (NTRS)

Kibbey, Timothy P.

2015-01-01

The ideal rocket equation is manipulated to demonstrate the essential link between propellant density and specific impulse as the two primary stage performance drivers for a launch vehicle. This is illustrated by examining volume-limited stages such as first stages and boosters. This proves to be a good approximation for first-order or Phase A vehicle design studies for solid rocket motors and for liquid stages, except when comparing to hydrogen-fueled stages. A next-order mass model is developed that is able to model the mass differences between hydrogen-fueled and other stages. Propellants considered range in density from liquid methane to inhibited red fuming nitric acid. Calculated comparisons are shown for solid rocket boosters, liquid first stages, liquid upper stages, and a balloon-deployed single-stage-to-orbit concept. The derived relationships are ripe for inclusion in a multi-stage design space exploration and optimization algorithm, as well as for single-parameter comparisons such as those shown herein.

Impulse noise trauma during army weapon firing.

PubMed

Munjal, K R; Singh, V P

1997-04-01

A 100 infanty personnel firing modern weapons such as the Anti Tank Guided Missile, 106mm Recoiless Gun (RCL), 84mm Rocket Launcher (RL) and 81mm Mortar were studied for the effect of impulse noise on the ear and the evolution of the Temporary Threshold Shift (TTS), Recovery Time (RT) and Permanent Threshold Shift (PTS) was traced.

Multivariable optimization of liquid rocket engines using particle swarm algorithms

NASA Astrophysics Data System (ADS)

Jones, Daniel Ray

Liquid rocket engines are highly reliable, controllable, and efficient compared to other conventional forms of rocket propulsion. As such, they have seen wide use in the space industry and have become the standard propulsion system for launch vehicles, orbit insertion, and orbital maneuvering. Though these systems are well understood, historical optimization techniques are often inadequate due to the highly non-linear nature of the engine performance problem. In this thesis, a Particle Swarm Optimization (PSO) variant was applied to maximize the specific impulse of a finite-area combustion chamber (FAC) equilibrium flow rocket performance model by controlling the engine's oxidizer-to-fuel ratio and de Laval nozzle expansion and contraction ratios. In addition to the PSO-controlled parameters, engine performance was calculated based on propellant chemistry, combustion chamber pressure, and ambient pressure, which are provided as inputs to the program. The performance code was validated by comparison with NASA's Chemical Equilibrium with Applications (CEA) and the commercially available Rocket Propulsion Analysis (RPA) tool. Similarly, the PSO algorithm was validated by comparison with brute-force optimization, which calculates all possible solutions and subsequently determines which is the optimum. Particle Swarm Optimization was shown to be an effective optimizer capable of quick and reliable convergence for complex functions of multiple non-linear variables.

The nuclear thermal electric rocket: a proposed innovative propulsion concept for manned interplanetary missions

NASA Astrophysics Data System (ADS)

Dujarric, C.; Santovincenzo, A.; Summerer, L.

2013-03-01

Conventional propulsion technology (chemical and electric) currently limits the possibilities for human space exploration to the neighborhood of the Earth. If farther destinations (such as Mars) are to be reached with humans on board, a more capable interplanetary transfer engine featuring high thrust, high specific impulse is required. The source of energy which could in principle best meet these engine requirements is nuclear thermal. However, the nuclear thermal rocket technology is not yet ready for flight application. The development of new materials which is necessary for the nuclear core will require further testing on ground of full-scale nuclear rocket engines. Such testing is a powerful inhibitor to the nuclear rocket development, as the risks of nuclear contamination of the environment cannot be entirely avoided with current concepts. Alongside already further matured activities in the field of space nuclear power sources for generating on-board power, a low level investigation on nuclear propulsion has been running since long within ESA, and innovative concepts have already been proposed at an IAF conference in 1999 [1, 2]. Following a slow maturation process, a new concept was defined which was submitted to a concurrent design exercise in ESTEC in 2007. Great care was taken in the selection of the design parameters to ensure that this quite innovative concept would in all respects likely be feasible with margins. However, a thorough feasibility demonstration will require a more detailed design including the selection of appropriate materials and the verification that these can withstand the expected mechanical, thermal, and chemical environment. So far, the predefinition work made clear that, based on conservative technology assumptions, a specific impulse of 920 s could be obtained with a thrust of 110 kN. Despite the heavy engine dry mass, a preliminary mission analysis using conservative assumptions showed that the concept was reducing the required Initial Mass in Low Earth Orbit compared to conventional nuclear thermal rockets for a human mission to Mars. Of course, the realization of this concept still requires proper engineering and the dimensioning of quite unconventional machinery. A patent was filed on the concept. Because of the operating parameters of the nuclear core, which are very specific to this type of concept, it seems possible to test on ground this kind of engine at full scale in close loop using a reasonable size test facility with safe and clean conditions. Such tests can be conducted within fully confined enclosure, which would substantially increase the associated inherent nuclear safety levels. This breakthrough removes a showstopper for nuclear rocket engines development. The present paper will disclose the NTER (Nuclear Thermal Electric Rocket) engine concept, will present some of the results of the ESTEC concurrent engineering exercise, and will explain the concept for the NTER on-ground testing facility. Regulations and safety issues related to the development and implementation of the NTER concept will be addressed as well.

Hydrodynamic Stability Analysis of Particle-Laden Solid Rocket Motors

NASA Astrophysics Data System (ADS)

Elliott, T. S.; Majdalani, J.

2014-11-01

Fluid-wall interactions within solid rocket motors can result in parietal vortex shedding giving rise to hydrodynamic instabilities, or unsteady waves, that translate into pressure oscillations. The oscillations can result in vibrations observed by the rocket, rocket subsystems, or payload, which can lead to changes in flight characteristics, design failure, or other undesirable effects. For many years particles have been embedded in solid rocket propellants with the understanding that their presence increases specific impulse and suppresses fluctuations in the flowfield. This study utilizes a two dimensional framework to understand and quantify the aforementioned two-phase flowfield inside a motor case with a cylindrical grain perforation. This is accomplished through the use of linearized Navier-Stokes equations with the Stokes drag equation and application of the biglobal ansatz. Obtaining the biglobal equations for analysis requires quantification of the mean flowfield within the solid rocket motor. To that end, the extended Taylor-Culick form will be utilized to represent the gaseous phase of the mean flowfield while the self-similar form will be employed for the particle phase. Advancing the mean flowfield by quantifying the particle mass concentration with a semi-analytical solution the finalized mean flowfield is combined with the biglobal equations resulting in a system of eight partial differential equations. This system is solved using an eigensolver within the framework yielding the entire spectrum of eigenvalues, frequency and growth rate components, at once. This work will detail the parametric analysis performed to demonstrate the stabilizing and destabilizing effects of particles within solid rocket combustion.

A Plasma Rocket Demonstration on the International Space Station

NASA Astrophysics Data System (ADS)

Petro, A.

2002-01-01

in the development of a magneto-plasma rocket for several years. This type of rocket could be used in the future to propel interplanetary spacecraft. One feature of this concept is the ability to vary its specific impulse so that it can be operated in a mode that maximizes propellant efficiency or a mode that maximizes thrust. For this reason the system is called the Variable Specific Impulse Magneto-plasma Rocket or VASIMR. This ability to vary specific impulse and thrust will allow for optimum low thrust interplanetary trajectories and results in shorter trip times than is possible with fixed specific impulse systems while preserving adequate payload margins. demonstrations are envisioned. A ground-based experiment of a low-power VASIMR prototype rocket is currently underway at the Advanced Space Propulsion Laboratory. The next step is a proposal to build and fly a 25-kilowatt VASIMR rocket as an external payload on the International Space Station. This experiment will provide an opportunity to demonstrate the performance of the rocket in space and measure the induced environment. The experiment will also utilize the space station for its intended purpose as a laboratory with vacuum conditions that cannot be matched by any laboratory on Earth. propulsion on the space station. An electric propulsion system like VASIMR, if provided with sufficient electrical power, could provide continuous drag force compensation for the space station. Drag compensation would eliminate the need for reboosting the station, an operation that will consume about 60 metric tons of propellant in a ten-year period. In contrast, an electric propulsion system would require very little propellant. In fact, a system like VASIMR can use waste hydrogen from the station's life support system as its propellant. This waste hydrogen is otherwise dumped overboard. Continuous drag compensation would also improve the microgravity conditions on the station. So electric propulsion can reduce propellant delivery requirements and thereby increase available payload capacity and at the same time improve the conditions for scientific research. and the space environment. This is a beneficial effect that prevents a charge buildup on the station. The station already operates two dedicated non-propulsive plasma contactor devices for this purpose. A VASIMR rocket would function as an additional plasma contactor. would be delivered to orbit in the Space Shuttle payload bay. It would be mounted on a standard payload attachment structure. After removal from the payload bay by the shuttle robotic arm, it would be handed to the space station robotic arm which would place it at an external payload attach site on the station truss. A mating device for power and data connections exists at the payload site. The experiment would receive one to three kilowatts of power from the station. About 600 watts would be used for cryogenic cooling and control devices. Additional power would be stored in a set of batteries. The VASIMR experiment would be operated for short periods when the batteries can provide power to the amplifiers that feed radio-frequency power to the thruster assembly. The thruster assembly is composed of an inner tube in which the neutral propellant is injected and ionized and a larger tube, which supports the radio frequency antennas, which ionize the gas and heat the plasma. Electromagnet coils that provide the magnetic field to constrain the flow of the plasma and form the magnetic exit nozzle surround these tubes. to this supply are planned for the experiment. The experiment will carry two dedicated propellant tanks which each have the capacity to store all the propellant needed for an experimental program lasting several months. With two propellant tanks, the opportunity exists to perform experiments with more than one type of propellant. Hydrogen is the primary choice for propellant but deuterium and helium are also of interest and might also be included. All the propellant is stored and used in gaseous form at ambient temperature. rocket. There is a superconducting electromagnet that will need to be maintained at cryogenic temperatures in order to operate properly. The magnet is in close proximity to the plasma so a combination of compact insulation and passive and active heat transport techniques will be employed. activity requirements. However, provisions will be included to capitalize on the presence of humans in case repairs or servicing is required. The batteries, propellant tanks, and electronic components will be designed for on-orbit removal and replacement, if necessary. could be located on the station to provide useful thrust for drag compensation. In order to provide power for continuous thrusting, it may be necessary to augment the power generation system for the station. Another attractive possibility is to develop an electric propulsion testbed for the space station. This testbed could be used for testing and certifying a variety of propulsion systems at various stages of maturity while providing thrust for the space station. This station facility would be a valuable asset for commercial and government space transportation programs. more powerful and capable propulsion systems that will be demonstrated on free-flying spacecraft in near-Earth space and eventually on missions to the planets.

Modeling and Testing of Non-Nuclear, Highpower Simulated Nuclear Thermal Rocket Reactor Elements

NASA Technical Reports Server (NTRS)

Kirk, Daniel R.

2005-01-01

When the President offered his new vision for space exploration in January of 2004, he said, "Our third goal is to return to the moon by 2020, as the launching point for missions beyond," and, "With the experience and knowledge gained on the moon, we will then be ready to take the next steps of space exploration: human missions to Mars and to worlds beyond." A human mission to Mars implies the need to move large payloads as rapidly as possible, in an efficient and cost-effective manner. Furthermore, with the scientific advancements possible with Project Prometheus and its Jupiter Icy Moons Orbiter (JIMO), (these use electric propulsion), there is a renewed interest in deep space exploration propulsion systems. According to many mission analyses, nuclear thermal propulsion (NTP), with its relatively high thrust and high specific impulse, is a serious candidate for such missions. Nuclear rockets utilize fission energy to heat a reactor core to very high temperatures. Hydrogen gas flowing through the core then becomes superheated and exits the engine at very high exhaust velocities. The combination of temperature and low molecular weight results in an engine with specific impulses above 900 seconds. This is almost twice the performance of the LOX/LH2 space shuttle engines, and the impact of this performance would be to reduce the trip time of a manned Mars mission from the 2.5 years, possible with chemical engines, to about 12-14 months.

Analysis of gas turbine engines using water and oxygen injection to achieve high Mach numbers and high thrust

NASA Technical Reports Server (NTRS)

Henneberry, Hugh M.; Snyder, Christopher A.

1993-01-01

An analysis of gas turbine engines using water and oxygen injection to enhance performance by increasing Mach number capability and by increasing thrust is described. The liquids are injected, either separately or together, into the subsonic diffuser ahead of the engine compressor. A turbojet engine and a mixed-flow turbofan engine (MFTF) are examined, and in pursuit of maximum thrust, both engines are fitted with afterburners. The results indicate that water injection alone can extend the performance envelope of both engine types by one and one-half Mach numbers at which point water-air ratios reach 17 or 18 percent and liquid specific impulse is reduced to some 390 to 470 seconds, a level about equal to the impulse of a high energy rocket engine. The envelope can be further extended, but only with increasing sacrifices in liquid specific impulse. Oxygen-airflow ratios as high as 15 percent were investigated for increasing thrust. Using 15 percent oxygen in combination with water injection at high supersonic Mach numbers resulted in thrust augmentation as high as 76 percent without any significant decrease in liquid specific impulse. The stoichiometric afterburner exit temperature increased with increasing oxygen flow, reaching 4822 deg R in the turbojet engine at a Mach number of 3.5. At the transonic Mach number of 0.95 where no water injection is needed, an oxygen-air ratio of 15 percent increased thrust by some 55 percent in both engines, along with a decrease in liquid specific impulse of 62 percent. Afterburner temperature was approximately 4700 deg R at this high thrust condition. Water and/or oxygen injection are simple and straightforward strategies to improve engine performance and they will add little to engine weight. However, if large Mach number and thrust increases are required, liquid flows become significant, so that operation at these conditions will necessarily be of short duration.

2011-2012 Dryden Center Innovation Fund End of the Year Report: Altitude-Compensating Rocket Nozzles

NASA Technical Reports Server (NTRS)

Jones, Daniel S.; Bui, Trong T.

2012-01-01

This report highlights one of the many successful projects at the NASA Dryden Flight Research Center that was approved for FY12 funding under the Center Innovation Fund. This project was focused on advancing the technology readiness level of one specific type of altitude-compensating nozzle: the dual-bell rocket nozzle. When considering a rocket's performance over its entire integrated trajectory, the dual-bell nozzle has been predicted to achieve a higher total impulse over the conventional bell nozzle, which is expected to result in a greater capability of payload mass to low-Earth orbit. Although the dual-bell rocket nozzle has been thoroughly studied for several decades, this nozzle has still not been adequately tested in a relevant flight-like environment. This report provides highlights and top-level details on the FY12 feasibility effort to advance this promising technology through flight test, a collaborative effort which leverages NASA Marshall's dual-bell nozzle research and development with Dryden's expertise in propulsion-focused flight testing. To accomplish this goal, the NASA F-15B is proposed as the testbed for the initial flight-test campaign to advance this greatly needed capability.

Experimental analysis of SiC-based refractory concrete in hybrid rocket nozzles

NASA Astrophysics Data System (ADS)

D'Elia, Raffaele; Bernhart, Gérard; Hijlkema, Jouke; Cutard, Thierry

2016-09-01

Hybrid propulsion represents a good alternative to the more widely used liquid and solid systems. This technology combines some important specifications of the latters, as the possibility of re-ignition, thrust modulation, a higher specific impulse than solid systems, a greater simplicity and a lower cost than liquid systems. Nevertheless the highly oxidizing environment represents a major problem as regards the thermo-oxidation and ablative behavior of nozzle materials. The main goal of this research is to characterize a silicon carbide based micro-concrete with a maximum aggregates size of 800 μm, in a hybrid propulsion environment. The nozzle throat has to resist to a highly oxidizing polyethylene/nitrous oxide hybrid environment, under temperatures up to 2900 K. Three tests were performed on concrete-based nozzles in HERA Hybrid Rocket Motor (HRM) test bench at ONERA. Pressure chamber evolution and observations before and after tests are used to investigate the ablated surface at nozzle throat. Ablation behavior and crack generation are discussed and some improvements are proposed.

A unique nuclear thermal rocket engine using a particle bed reactor

NASA Astrophysics Data System (ADS)

Culver, Donald W.; Dahl, Wayne B.; McIlwain, Melvin C.

1992-01-01

Aerojet Propulsion Division (APD) studied 75-klb thrust Nuclear Thermal Rocket Engines (NTRE) with particle bed reactors (PBR) for application to NASA's manned Mars mission and prepared a conceptual design description of a unique engine that best satisfied mission-defined propulsion requirements and customer criteria. This paper describes the selection of a sprint-type Mars transfer mission and its impact on propulsion system design and operation. It shows how our NTRE concept was developed from this information. The resulting, unusual engine design is short, lightweight, and capable of high specific impulse operation, all factors that decrease Earth to orbit launch costs. Many unusual features of the NTRE are discussed, including nozzle area ratio variation and nozzle closure for closed loop after cooling. Mission performance calculations reveal that other well known engine options do not support this mission.

Flight and Preflight Tests of a Ram Jet Burning Magnesium Slurry Fuel and Utilizing a Solid-propellant Gas Generator for Fuel Expulsion

NASA Technical Reports Server (NTRS)

Bartlett, Walter, A , jr; Hagginbotham, William K , Jr

1955-01-01

Data obtained from the first flight test of a ram jet utilizing a magnesium slurry fuel are presented. The ram jet accelerated from a Mach number of 1.75 to a Mach number of 3.48 in 15.5 seconds. During this period a maximum values of air specific impulse and gross thrust coefficient were calculated to be 151 seconds and 0.658, respectively. The rocket gas generator used as a fuel-pumping system operated successfully.

Hybrid rocket propulsion systems for outer planet exploration missions

NASA Astrophysics Data System (ADS)

Jens, Elizabeth T.; Cantwell, Brian J.; Hubbard, G. Scott

2016-11-01

Outer planet exploration missions require significant propulsive capability, particularly to achieve orbit insertion. Missions to explore the moons of outer planets place even more demanding requirements on propulsion systems, since they involve multiple large ΔV maneuvers. Hybrid rockets present a favorable alternative to conventional propulsion systems for many of these missions. They typically enjoy higher specific impulse than solids, can be throttled, stopped/restarted, and have more flexibility in their packaging configuration. Hybrids are more compact and easier to throttle than liquids and have similar performance levels. In order to investigate the suitability of these propulsion systems for exploration missions, this paper presents novel hybrid motor designs for two interplanetary missions. Hybrid propulsion systems for missions to Europa and Uranus are presented and compared to conventional in-space propulsion systems. The hybrid motor design for each of these missions is optimized across a range of parameters, including propellant selection, O/F ratio, nozzle area ratio, and chamber pressure. Details of the design process are described in order to provide guidance for researchers wishing to evaluate hybrid rocket motor designs for other missions and applications.

High Power Ion Cyclotron Heating in the VASIMR

NASA Astrophysics Data System (ADS)

Longmier, B. W.; Brukardt, M. S.; Bering, E. A.; Chang Diaz, F.; Squire, J.

2009-12-01

The Variable Specific Impulse Magnetoplasma Rocket (VASIMR®) is an electric propulsion system under development at Ad Astra Rocket Company that utilizes several processes of ion acceleration and heating that occur in the Birkeland currents of an auroral arc system. Among these processes are parallel electric field acceleration, lower hybrid resonance heating, and ion cyclotron resonance heating. The VASIMR® is capable of laboratory simulation of electromagnetic ion cyclotron wave heating during a single pass of ions through the resonance region. The plasma is generated by a helicon discharge of 35 kW then passes through a 176 kW RF booster stage that couples left hand polarized slow mode waves from the high field side of the resonance. VX-200 auroral simulation results from the past year are discussed. Ambipolar acceleration has been shown to produce 35eV argon ions in the helicon exhaust. The effects on the ion exhaust with an addition of 150-200 kW of ion cyclotron heating are presented. The changes to the VASIMR® experiment at Ad Astra Rocket Company's new facility in Webster, Texas will also be discussed, including the possibility of collaborative experiments.

Ion Cyclotron Waves in the VASIMR

NASA Astrophysics Data System (ADS)

Brukardt, M. S.; Bering, E. A.; Chang-Diaz, F. R.; Squire, J. P.; Longmier, B.

2008-12-01

The Variable Specific Impulse Magnetoplasma Rocket is an electric propulsion system under development at Ad Astra Rocket Company that utilizes several processes of ion acceleration and heating that occur in the Birkeland currents of an auroral arc system. Among these processes are parallel electric field acceleration, lower hybrid resonance heating, and ion cyclotron resonance heating. The VASIMR is capable of laboratory simulation of electromagnetic ion cyclotron wave heating during a single pass of the plasma through the resonance region. The plasma is generated by a helicon discharge of about 25 kW then passes through an RF booster stage that shoots left hand polarized slow mode waves from the high field side of the resonance. This paper will focus on the upgrades to the VX-200 test model over the last year. After summarizing the VX- 50 and VX-100 results, the new data from the VX-200 model will be presented. Lastly, the changes to the VASIMR experiment due to Ad Astra Rocket Company's new facility in Webster, Texas will also be discussed, including the possibility of collaborative experiments at the new facility.

Thrust augmentation nozzle (TAN) concept for rocket engine booster applications

NASA Astrophysics Data System (ADS)

Forde, Scott; Bulman, Mel; Neill, Todd

2006-07-01

Aerojet used the patented thrust augmented nozzle (TAN) concept to validate a unique means of increasing sea-level thrust in a liquid rocket booster engine. We have used knowledge gained from hypersonic Scramjet research to inject propellants into the supersonic region of the rocket engine nozzle to significantly increase sea-level thrust without significantly impacting specific impulse. The TAN concept overcomes conventional engine limitations by injecting propellants and combusting in an annular region in the divergent section of the nozzle. This injection of propellants at moderate pressures allows for obtaining high thrust at takeoff without overexpansion thrust losses. The main chamber is operated at a constant pressure while maintaining a constant head rise and flow rate of the main propellant pumps. Recent hot-fire tests have validated the design approach and thrust augmentation ratios. Calculations of nozzle performance and wall pressures were made using computational fluid dynamics analyses with and without thrust augmentation flow, resulting in good agreement between calculated and measured quantities including augmentation thrust. This paper describes the TAN concept, the test setup, test results, and calculation results.

Rocket propulsion by thermonuclear micro-bombs ignited with intense relativistic electron beams.

NASA Technical Reports Server (NTRS)

Winterberg, F.

1971-01-01

Discussion of a method for the ignition of a thermonuclear microbomb by means of an intense relativistic electron beam with regard to its potential application to rocket propulsion. With such a system, exhaust velocities up to 1000 km/sec, corresponding to a specific impulse of 100,000 sec, seem to be within the realm of possibility. The rocket is propelled by a chain of thermonuclear microbombs exploded in a concave magnetic mirror produced by superconducting field coils. The magnetic pressure of the field reflects the fireball generated by the explosion. For the large capacitor bank required to generate the intense relativistic electron beam, a desirable lightweight design may be possible through use of ferroelectric materials. Because of the high cost of the T-D and He 3-D thermonuclear material, the system has to be optimized by minimizing the T-D and He 3-D consumption by a proper TD and He 3-D fuel to hydrogen propellant mass ratio, leading to a larger total system mass than would be absolutely necessary.

Nuclear Thermal Rocket Simulation in NPSS

NASA Technical Reports Server (NTRS)

Belair, Michael L.; Sarmiento, Charles J.; Lavelle, Thomas M.

2013-01-01

Four nuclear thermal rocket (NTR) models have been created in the Numerical Propulsion System Simulation (NPSS) framework. The models are divided into two categories. One set is based upon the ZrC-graphite composite fuel element and tie tube-style reactor developed during the Nuclear Engine for Rocket Vehicle Application (NERVA) project in the late 1960s and early 1970s. The other reactor set is based upon a W-UO2 ceramic-metallic (CERMET) fuel element. Within each category, a small and a large thrust engine are modeled. The small engine models utilize RL-10 turbomachinery performance maps and have a thrust of approximately 33.4 kN (7,500 lbf ). The large engine models utilize scaled RL-60 turbomachinery performance maps and have a thrust of approximately 111.2 kN (25,000 lbf ). Power deposition profiles for each reactor were obtained from a detailed Monte Carlo N-Particle (MCNP5) model of the reactor cores. Performance factors such as thermodynamic state points, thrust, specific impulse, reactor power level, and maximum fuel temperature are analyzed for each engine design.

Nuclear Thermal Rocket Simulation in NPSS

NASA Technical Reports Server (NTRS)

Belair, Michael L.; Sarmiento, Charles J.; Lavelle, Thomas L.

2013-01-01

Four nuclear thermal rocket (NTR) models have been created in the Numerical Propulsion System Simulation (NPSS) framework. The models are divided into two categories. One set is based upon the ZrC-graphite composite fuel element and tie tube-style reactor developed during the Nuclear Engine for Rocket Vehicle Application (NERVA) project in the late 1960s and early 1970s. The other reactor set is based upon a W-UO2 ceramic- metallic (CERMET) fuel element. Within each category, a small and a large thrust engine are modeled. The small engine models utilize RL-10 turbomachinery performance maps and have a thrust of approximately 33.4 kN (7,500 lbf ). The large engine models utilize scaled RL-60 turbomachinery performance maps and have a thrust of approximately 111.2 kN (25,000 lbf ). Power deposition profiles for each reactor were obtained from a detailed Monte Carlo N-Particle (MCNP5) model of the reactor cores. Performance factors such as thermodynamic state points, thrust, specific impulse, reactor power level, and maximum fuel temperature are analyzed for each engine design.

Pressure oscillations and instability of working processes in the combustion chambers of solid rocket motors

NASA Astrophysics Data System (ADS)

Emelyanov, V. N.; Teterina, I. V.; Volkov, K. N.; Garkushev, A. U.

2017-06-01

Metal particles are widely used in space engineering to increase specific impulse and to supress acoustic instability of intra-champber processes. A numerical analysis of the internal injection-driven turbulent gas-particle flows is performed to improve the current understanding and modeling capabilities of the complex flow characteristics in the combustion chambers of solid rocket motors (SRMs) in presence of forced pressure oscillations. The two-phase flow is simulated with a combined Eulerian-Lagrangian approach. The Reynolds-averaged Navier-Stokes equations and transport equations of k - ε model are solved numerically for the gas. The particulate phase is simulated through a Lagrangian deterministic and stochastic tracking models to provide particle trajectories and particle concentration. The results obtained highlight the crucial significance of the particle dispersion in turbulent flowfield and high potential of statistical methods. Strong coupling between acoustic oscillations, vortical motion, turbulent fluctuations and particle dynamics is observed.

A Collimated Retarding Potential Analyzer for the Study of Magnetoplasma Rocket Plumes

NASA Technical Reports Server (NTRS)

Glover, T. W.; Chan, A. A.; Chang-Diaz, F. R.; Kittrell, C.

2003-01-01

A gridded retarding potential analyzer (RPA) has been developed to characterize the magnetized plasma exhaust of the 10 kW Variable Specific Impulse Magnetoplasma Rocket (VX-10) experiment at NASA's Advanced Space Propulsion Laboratory. In this system, plasma is energized through coupling of radio frequency waves at the ion cyclotron resonance (ICR). The particles are subsequently accelerated in a magnetic nozzle to provide thrust. Downstream of the nozzle, the RPA's mounting assembly enables the detector to make complete axial and radial scans of the plasma. A multichannel collimator can be inserted into the RPA to remove ions with pitch angles greater than approximately 1 deg. A calculation of the general collimator transmission as a function over velocity space is presented, which shows the instrument's sensitivity in detecting changes in both the parallel and perpendicular components of the ion energy. Data from initial VX-10 ICRH experiments show evidence of ion heating.

The engineering of a nuclear thermal landing and ascent vehicle utilizing indigenous Martian propellant

NASA Technical Reports Server (NTRS)

Zubrin, Robert M.

1990-01-01

A design study of a novel space transportation concept called NIMF (Nuclear rocket using Indigenous Martian Fuel) is reported. In this concept, Martian CO2 gas, which constitutes 95 percent of the atmosphere, is liquified by simple compression to about 100 psi and remains stable without refrigeration. When heated and exhausted out of a rocket nozzle, a specific impulse of about 264 s can be achieved, sufficient for flights from the surface to highly energetic orbits or from one point on the surface to any other point. The propellant acquisition system can travel with the vehicle, allowing it to refuel itself each time it lands. The concept offers unequalled potential to achieve planetwide mobility, allowing complete global access for the exploration of Mars. By eliminating the necessity of transporting ascent propellant to Mars, the NIMF can also significantly reduce the initial mass in LEO and of a manned Mars mission.

Estimating Basic Preliminary Design Performances of Aerospace Vehicles

NASA Technical Reports Server (NTRS)

Luz, Paul L.; Alexander, Reginald

2004-01-01

Aerodynamics and Performance Estimation Toolset is a collection of four software programs for rapidly estimating the preliminary design performance of aerospace vehicles represented by doing simplified calculations based on ballistic trajectories, the ideal rocket equation, and supersonic wedges through standard atmosphere. The program consists of a set of Microsoft Excel worksheet subprograms. The input and output data are presented in a user-friendly format, and calculations are performed rapidly enough that the user can iterate among different trajectories and/or shapes to perform "what-if" studies. Estimates that can be computed by these programs include: 1. Ballistic trajectories as a function of departure angles, initial velocities, initial positions, and target altitudes; assuming point masses and no atmosphere. The program plots the trajectory in two-dimensions and outputs the position, pitch, and velocity along the trajectory. 2. The "Rocket Equation" program calculates and plots the trade space for a vehicle s propellant mass fraction over a range of specific impulse and mission velocity values, propellant mass fractions as functions of specific impulses and velocities. 3. "Standard Atmosphere" will estimate the temperature, speed of sound, pressure, and air density as a function of altitude in a standard atmosphere, properties of a standard atmosphere as functions of altitude. 4. "Supersonic Wedges" will calculate the free-stream, normal-shock, oblique-shock, and isentropic flow properties for a wedge-shaped body flying supersonically through a standard atmosphere. It will also calculate the maximum angle for which a shock remains attached, and the minimum Mach number for which a shock becomes attached, all as functions of the wedge angle, altitude, and Mach number.

Performance Increase Verification for a Bipropellant Rocket Engine

NASA Technical Reports Server (NTRS)

Alexander, Leslie; Chapman, Jack; Wilson, Reed; Krismer, David; Lu, Frank; Wilson, Kim; Miller, Scott; England, Chris

2008-01-01

Component performance assessment testing for a, pressure-fed earth storable bipropellant rocket engine was successfully completed at Aerojet's Redmond test facility. The primary goal of the this development project is to increase the specific impulse of an apogee class bi-propellant engine to greater than 330 seconds with nitrogen tetroxide and monomethylhydrazine propellants and greater than 335 seconds with nitrogen tetroxide and hydrazine. The secondary goal of the project is to take greater advantage of the high temperature capabilities of iridium/rhenium chambers. In order to achieve these goals, the propellant feed pressures were increased to 400 psia, nominal, which in turn increased the chamber pressure and temperature, allowing for higher c*. The tests article used a 24-on-24 unlike doublet injector design coupled with a copper heat sink chamber to simulate a flight configuration combustion chamber. The injector is designed to produce a nominal 200 lbf of thrust with a specific impulse of 335 seconds (using hydrazine fuel). Effect of Chamber length on engine C* performance was evaluated with the use of modular, bolt-together test hardware and removable chamber inserts. Multiple short duration firings were performed to characterize injector performance across a range of thrust levels, 180 to 220 lbf, and mixture ratios, from 1.1 to 1.3. During firing, ignition transient, chamber pressure, and various temperatures were measured in order to evaluate the performance of the engine and characterize the thermal conditions. The tests successfully demonstrated the stable operation and performance potential of a full scale engine with a measured c* of XXXX ft/sec (XXXX m/s) under nominal operational conditions.

Concept and performance study of turbocharged solid propellant ramjet

NASA Astrophysics Data System (ADS)

Li, Jiang; Liu, Kai; Liu, Yang; Liu, Shichang

2018-06-01

This study proposes a turbocharged solid propellant ramjet (TSPR) propulsion system that integrates a turbocharged system consisting of a solid propellant (SP) air turbo rocket (ATR) and the fuel-rich gas generator of a solid propellant ramjet (SPR). First, a suitable propellant scheme was determined for the TSPR. A solid hydrocarbon propellant is used to generate gas for driving the turbine, and a boron-based fuel-rich propellant is used to provide fuel-rich gas to the afterburner. An appropriate TSPR structure was also determined. The TSPR's thermodynamic cycle was analysed to prove its theoretical feasibility. The results showed that the TSPR's specific cycle power was larger than those of SP-ATR and SPR and thermal efficiency was slightly less than that of SP-ATR. Overall, TSPR showed optimal performance in a wide flight envelope. The specific impulses and specific thrusts of TSPR, SP-ATR, and SPR in the flight envelope were calculated and compared. TSPR's flight envelope roughly overlapped that of SP-ATR, its specific impulse was larger than that of SP-ATR, and its specific thrust was larger than those of SP-ATR and SPR. Attempts to improve the TSPR off-design performance prompted our proposal of a control plan for off-design codes in which both the turbocharger corrected speed and combustor excess gas coefficient are kept constant. An off-design performance model was established by analysing the TSPR working process. We concluded that TSPR with a constant corrected speed had wider flight envelope, higher thrust, and higher specific impulse than TSPR with a constant physical speed determined by calculating the performance of off-design TSPR codes under different control plans. The results of this study can provide a reference for further studies on TSPRs.

High-Performance Bipropellant Engine

NASA Technical Reports Server (NTRS)

Biaglow, James A.; Schneider, Steven J.

1999-01-01

TRW, under contract to the NASA Lewis Research Center, has successfully completed over 10 000 sec of testing of a rhenium thrust chamber manufactured via a new-generation powder metallurgy. High performance was achieved for two different propellants, N2O4- N2H4 and N2O4 -MMH. TRW conducted 44 tests with N2O4-N2H4, accumulating 5230 sec of operating time with maximum burn times of 600 sec and a specific impulse Isp of 333 sec. Seventeen tests were conducted with N2O4-MMH for an additional 4789 sec and a maximum Isp of 324 sec, with a maximum firing duration of 700 sec. Together, the 61 tests totalled 10 019 sec of operating time, with the chamber remaining in excellent condition. Of these tests, 11 lasted 600 to 700 sec. The performance of radiation-cooled rocket engines is limited by their operating temperature. For the past two to three decades, the majority of radiation-cooled rockets were composed of a high-temperature niobium alloy (C103) with a disilicide oxide coating (R512) for oxidation resistance. The R512 coating practically limits the operating temperature to 1370 C. For the Earth-storable bipropellants commonly used in satellite and spacecraft propulsion systems, a significant amount of fuel film cooling is needed. The large film-cooling requirement extracts a large penalty in performance from incomplete mixing and combustion. A material system with a higher temperature capability has been matured to the point where engines are being readied for flight, particularly the 100-lb-thrust class engine. This system has powder rhenium (Re) as a substrate material with an iridium (Ir) oxidation-resistant coating. Again, the operating temperature is limited by the coating; however, Ir is capable of long-life operation at 2200 C. For Earth-storable bipropellants, this allows for the virtual elimination of fuel film cooling (some film cooling is used for thermal control of the head end). This has resulted in significant increases in specific impulse performance (15 to 20 sec). To determine the merits of a powder rhenium thrust chamber, Lewis On-Board Propulsion Branch directed TRW (under the Space Storable Rocket Technology Program and the High Pressure Earth Storable Rocket Technology Program) to design, fabricate, and test an engineering model to serve as a technology demonstrator.

Small Launch Vehicle Concept Development for Affordable Multi-Stage Inline Configurations

NASA Technical Reports Server (NTRS)

Beers, Benjamin R.; Waters, Eric D.; Philips, Alan D.; Threet, Grady E., Jr.

2014-01-01

The Advanced Concepts Office at NASA's George C. Marshall Space Flight Center conducted a study of two configurations of a three stage, inline, liquid propellant small launch vehicle concept developed on the premise of maximizing affordability by targeting a specific payload capability range based on current industry demand. The initial configuration, NESC-1, employed liquid oxygen as the oxidizer and rocket propellant grade kerosene as the fuel in all three stages. The second and more heavily studied configuration, NESC-4, employed liquid oxygen and rocket propellant grade kerosene on the first and second stages and liquid oxygen and liquid methane fuel on the third stage. On both vehicles, sensitivity studies were first conducted on specific impulse and stage propellant mass fraction in order to baseline gear ratios and drive the focus of concept development. Subsequent sensitivity and trade studies on the NESC-4 configuration investigated potential impacts to affordability due to changes in gross liftoff weight and/or vehicle complexity. Results are discussed at a high level to understand the severity of certain sensitivities and how those trade studies conducted can either affect cost, performance or both.

Studies of the thermal and optical responses of H atoms in solid H2

NASA Technical Reports Server (NTRS)

Gaines, James R.; Vause, Chester A., III

1990-01-01

It was the goal of this reserch project to model both the storage of energy in solid hydrogen in the form of atoms and the conversion of this stored energy into other forms of useful energy. The basic ideas of rocket propulsion originate in classical physics and they remain unchanged. To escape a strong gravitational field, the 'burn time' must be minimized but in negligible force fields, the burn time is unimportant and only the relative masses of rocket to fuel determine a specific exhaust velocity. It is in this latter case that low mass fuels such as hydrogen become very important. The burning of hydrogen in oxygen is a 'benchmark' fuel today providing a specific impulse of 400 seconds or better. More exotic fuels will be needed for many of the interesting explorations of the future but they still must have large energy releases per unit mass. It is in this context that propulsion based on hydrogen atom recombination receives attention and these studies will serve as engineering guides.

Flight motor set 360L001 (STS-26R). Volume 1: System overview, revision A

NASA Technical Reports Server (NTRS)

Garecht, Diane M.

1990-01-01

The NASA space shuttle flight STS-26R, launched at 11:37.00.009 am, EDT on 29 Sep. 1988, used the redesigned solid rocket motors (RSRM) 360L001A and 360L001B. Evaluation of the ground environment instrumentation (GEI) data recorded prior to flight showed no launch commit criteria violations; that the field joint heater and aft skirt thermal conditioning systems performed adequately; and that the GEI data showed good agreement with thermal model predictions. Evaluation of the developmental flight instrumentation (DFI) revealed excellent agreement with both the predicted and required ballistic specifications. All parameters were well within the GEI specification requirements including propellant burn rates, specific impulse values, and thrust imbalance. Recorded strain values also indicated satisfactory radial growth and stress levels, as well as verification of adequate safety factors. Postflight inspection of the insulation, seals, case, and nozzles showed overall excellent performance. Some thermal DFI protective cork was missing, and inoperative field joint vent valves on the thermal protection cork allowed water entry into the field joints upon splashdown. Evaluation of these anomalies, as well as complete evaluation of all Redesigned Solid Rocket Motor components, is contained.

Exhaust Plume Measurements of the VASIMR VX-200

NASA Astrophysics Data System (ADS)

Longmier, Benjamin; Bering, Edgar, III; Squire, Jared; Glover, Tim; Chang-Diaz, Franklin; Brukardt, Michael

2008-11-01

Recent progress is discussed in the development of an advanced RF electric propulsion concept: the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) VX-200 engine, a 200 kW flight-technology prototype. Results from high power Helicon only and Helicon with ICRH experiments are performed on the VX-200 using argon plasma. Recent measurements of axial plasma density and potential profiles, magnetic field-line shaping, charge exchange, and force measurements taken in the plume of the VX-200 exhaust are made within a new 125 cubic meter cryo-pumped vacuum chamber and are presented in the context of RF plasma thruster physics.

Space shuttle main engine definition (phase B). Volume 2: Avionics. [for space shuttle

NASA Technical Reports Server (NTRS)

1971-01-01

The advent of the space shuttle engine with its requirements for high specific impulse, long life, and low cost have dictated a combustion cycle and a closed loop control system to allow the engine components to run close to operating limits. These performance requirements, combined with the necessity for low operational costs, have placed new demands on rocket engine control, system checkout, and diagnosis technology. Based on considerations of precision environment, and compatibility with vehicle interface commands, an electronic control, makes available many functions that logically provide the information required for engine system checkout and diagnosis.

Fusion Propulsion and Power for Future Flight

NASA Technical Reports Server (NTRS)

Froning, H. D., Jr.

1996-01-01

There are innovative magnetic and electric confinement fusion power and propulsion system designs with potential for: vacuum specific impulses of 1500-2000 seconds with rocket engine thrust/mass ratios of 5-10 g's; environmentally favorable exhaust emissions if aneutronic fusion propellants can be used; a 2 to 3-fold reduction in the mass of hypersonic airliners and SSTO aerospace planes; a 10 to 20 fold reduction in Mars expedition mass and cost (if propellant from planetary atmospheres is used); and feasibility or in-feasibility of these systems could be confirmed with a modest applied research and exploratory development cost.

Options For Development of Space Fission Propulsion Systems

NASA Technical Reports Server (NTRS)

Houta, Mike; VanDyke, Melissa; Godfroy, Tom; Pedersen, Kevin; Martin, James; Dickens, Ricky; Salvail, Pat; Hrbud, Ivana; Rodgers, Stephen L. (Technical Monitor)

2001-01-01

Fission technology can enable rapid, affordable access to any point in the solar system. Potential fission-based transportation options include high specific power continuous impulse propulsion systems and bimodal nuclear thermal rockets. Despite their tremendous potential for enhancing or enabling deep space and planetary missions, to date space fission system have only been used in Earth orbit. The first step towards utilizing advanced fission propulsion systems is development of a safe, near-term, affordable fission system that can enhance or enable near-term missions of interest. An evolutionary approach for developing space fission propulsion systems is proposed.

Comparison of Engine Cycle Codes for Rocket-Based Combined Cycle Engines

NASA Technical Reports Server (NTRS)

Waltrup, Paul J.; Auslender, Aaron H.; Bradford, John E.; Carreiro, Louis R.; Gettinger, Christopher; Komar, D. R.; McDonald, J.; Snyder, Christopher A.

2002-01-01

This paper summarizes the results from a one day workshop on Rocket-Based Combined Cycle (RBCC) Engine Cycle Codes held in Monterey CA in November of 2000 at the 2000 JANNAF JPM with the authors as primary participants. The objectives of the workshop were to discuss and compare the merits of existing Rocket-Based Combined Cycle (RBCC) engine cycle codes being used by government and industry to predict RBCC engine performance and interpret experimental results. These merits included physical and chemical modeling, accuracy and user friendliness. The ultimate purpose of the workshop was to identify the best codes for analyzing RBCC engines and to document any potential shortcomings, not to demonstrate the merits or deficiencies of any particular engine design. Five cases representative of the operating regimes of typical RBCC engines were used as the basis of these comparisons. These included Mach 0 sea level static and Mach 1.0 and Mach 2.5 Air-Augmented-Rocket (AAR), Mach 4 subsonic combustion ramjet or dual-mode scramjet, and Mach 8 scramjet operating modes. Specification of a generic RBCC engine geometry and concomitant component operating efficiencies, bypass ratios, fuel/oxidizer/air equivalence ratios and flight dynamic pressures were provided. The engine included an air inlet, isolator duct, axial rocket motor/injector, axial wall fuel injectors, diverging combustor, and exit nozzle. Gaseous hydrogen was used as the fuel with the rocket portion of the system using a gaseous H2/O2 propellant system to avoid cryogenic issues. The results of the workshop, even after post-workshop adjudication of differences, were surprising. They showed that the codes predicted essentially the same performance at the Mach 0 and I conditions, but progressively diverged from a common value (for example, for fuel specific impulse, Isp) as the flight Mach number increased, with the largest differences at Mach 8. The example cases and results are compared and discussed in this paper.

Performance Criteria of Nuclear Space Propulsion Systems

NASA Astrophysics Data System (ADS)

Shepherd, L. R.

Future exploration of the solar system on a major scale will require propulsion systems capable of performance far greater than is achievable with the present generation of rocket engines using chemical propellants. Viable missions going deeper into interstellar space will be even more demanding. Propulsion systems based on nuclear energy sources, fission or (eventually) fusion offer the best prospect for meeting the requirements. The most obvious gain coming from the application of nuclear reactions is the possibility, at least in principle, of obtaining specific impulses a thousandfold greater than can be achieved in chemically energised rockets. However, practical considerations preclude the possibility of exploiting the full potential of nuclear energy sources in any engines conceivable in terms of presently known technology. Achievable propulsive power is a particularly limiting factor, since this determines the acceleration that may be obtained. Conventional chemical rocket engines have specific propulsive powers (power per unit engine mass) in the order of gigawatts per tonne. One cannot envisage the possibility of approaching such a level of performance by orders of magnitude in presently conceivable nuclear propulsive systems. The time taken, under power, to reach a given terminal velocity is proportional to the square of the engine's exhaust velocity and the inverse of its specific power. An assessment of various nuclear propulsion concepts suggests that, even with the most optimistic assumptions, it could take many hundreds of years to attain the velocities necessary to reach the nearest stars. Exploration within a range of the order of a thousand AU, however, would appear to offer viable prospects, even with the low levels of specific power of presently conceivable nuclear engines.

Convective Heat Transfer in the Reusable Solid Rocket Motor of the Space Transportation System

NASA Technical Reports Server (NTRS)

Ahmad, Rashid A.; Cash, Stephen F. (Technical Monitor)

2002-01-01

This simulation involved a two-dimensional axisymmetric model of a full motor initial grain of the Reusable Solid Rocket Motor (RSRM) of the Space Transportation System (STS). It was conducted with CFD (computational fluid dynamics) commercial code FLUENT. This analysis was performed to: a) maintain continuity with most related previous analyses, b) serve as a non-vectored baseline for any three-dimensional vectored nozzles, c) provide a relatively simple application and checkout for various CFD solution schemes, grid sensitivity studies, turbulence modeling and heat transfer, and d) calculate nozzle convective heat transfer coefficients. The accuracy of the present results and the selection of the numerical schemes and turbulence models were based on matching the rocket ballistic predictions of mass flow rate, head end pressure, vacuum thrust and specific impulse, and measured chamber pressure drop. Matching these ballistic predictions was found to be good. This study was limited to convective heat transfer and the results compared favorably with existing theory. On the other hand, qualitative comparison with backed-out data of the ratio of the convective heat transfer coefficient to the specific heat at constant pressure was made in a relative manner. This backed-out data was devised to match nozzle erosion that was a result of heat transfer (convective, radiative and conductive), chemical (transpirating), and mechanical (shear and particle impingement forces) effects combined.

UEDGE Simulations for Power and Particle Flow Analysis of FRC Rocket

NASA Astrophysics Data System (ADS)

Zheng, Fred; Evans, Eugene S.; McGreivy, Nick; Kaptanoglu, Alan; Izacard, Olivier; Cohen, Samuel A.

2017-10-01

The field-reversed configuration (FRC) is under consideration for use in a direct fusion drive (DFD) rocket propulsion system for future space missions. To achieve a rocket configuration, the FRC is embedded within an asymmetric magnetic mirror, in which one end is closed and contains a gas box, and the other end is open and incorporates a magnetic nozzle. Neutral deuterium is injected into the gas box, and flows through the scrape-off layer (SOL) around the core plasma and out the magnetic nozzle, both cooling the core and serving as propellant. Previous studies have examined a range of operating conditions for the SOL of a DFD using UEDGE, a 2D fluid code; discrepancies on the order of 5% were found during the analysis of overall power balance. This work extends the analysis of the previously-studied SOL geometry by updating boundary conditions and conducting a detailed study of power and particle flows within the simulation with the goals of modeling electrical power generation instead of thrust and achieving higher specific impulse. This work was supported, in part, by DOE Contract Number DE-AC02-09CH11466 and Princeton Environmental Institute.

Fabrication of High Thermal Conductivity NARloy-Z-Diamond Composite Combustion Chamber Liner for Advanced Rocket Engines

NASA Technical Reports Server (NTRS)

Bhat, Biliyar N.; Greene, Sandra E.; Singh, Jogender

2016-01-01

This paper describes the process development for fabricating a high thermal conductivity NARloy-Z-Diamond composite (NARloy-Z-D) combustion chamber liner for application in advanced rocket engines. The fabrication process is challenging and this paper presents some details of these challenges and approaches used to address them. Prior research conducted at NASA-MSFC and Penn State had shown that NARloy-Z-40%D composite material has significantly higher thermal conductivity than the state of the art NARloy-Z alloy. Furthermore, NARloy-Z-40 %D is much lighter than NARloy-Z. These attributes help to improve the performance of the advanced rocket engines. Increased thermal conductivity will directly translate into increased turbopump power, increased chamber pressure for improved thrust and specific impulse. Early work on NARloy-Z-D composites used the Field Assisted Sintering Technology (FAST, Ref. 1, 2) for fabricating discs. NARloy-Z-D composites containing 10, 20 and 40vol% of high thermal conductivity diamond powder were investigated. Thermal conductivity (TC) data. TC increased with increasing diamond content and showed 50% improvement over pure copper at 40vol% diamond. This composition was selected for fabricating the combustion chamber liner using the FAST technique.

Feasibility of Colliding-beam fast-fission reactor via 238U80++238 U80+ --> 4 FF + 5n + 430 MeV beam with suppressed plutonium and direct conversion of fission fragment (FF) energy into electricity and/or Rocket propellant with high specific impulse

NASA Astrophysics Data System (ADS)

Maglich, Bogdan; Hester, Tim; Calsec Collaboration

2015-10-01

Uranium-uranium colliding beam experiment1, used fully ionized 238U92+ at energy 100GeV --> <-- 100 GeV, has measured total σ = 487 b. Reaction rate of colliding beams is proportional to neutron flux-squared. First functional Auto-Collider3-6, a compact Migma IV, 1 m in diameter, had self-colliding deuterons, D+, of 725 KeV --> <-- 725 KeV, resulting in copious production of T and 3He. U +U Autocollider``EXYDER'' will use strong-focusing magnet7, which would increase reaction rate by 104. 80 times ionized U ions accelerated through 3 MV accelerator, will collide beam 240 MeV --> <-- 240 MeV. Reaction is: 238U80+ +238 U80+ --> 4 FF + 5n + 430 MeV. Using a simple model1 fission σf ~ 100 b. Suppression of Pu by a factor of 106 will be achieved because NO thermal neutron fission can take place; only fast, 1-3 MeV, where σabs is negligible. Direct conversion of 95% of 430 MeV produced is carried by electrically charged FFs which are magnetically funneled for direct conversion of energy of FFs via electrostatic decelerators4,11. 90% of 930 MeV is electrically recoverable. Depending on the assumptions, we project electric _ power density production of 20 to 200 MWe m-3, equivalent to Thermal 1.3 - 13 GWthm-3. If one-half of unburned U is used for propulsion while rest powers system, heavy FF ion mass provides specific impulse Isp = 106 sec., 103 times higher than current rocket engines.

Development priorities for in-space propulsion technologies

NASA Astrophysics Data System (ADS)

Johnson, Les; Meyer, Michael; Palaszewski, Bryan; Coote, David; Goebel, Dan; White, Harold

2013-02-01

During the summer of 2010, NASA's Office of Chief Technologist assembled 15 civil service teams to support the creation of a NASA integrated technology roadmap. The Aero-Space Technology Area Roadmap is an integrated set of technology area roadmaps recommending the overall technology investment strategy and prioritization for NASA's technology programs. The integrated set of roadmaps will provide technology paths needed to meet NASA's strategic goals. The roadmaps have been reviewed by senior NASA management and the National Research Council. With the exception of electric propulsion systems used for commercial communications satellite station-keeping and a handful of deep space science missions, almost all of the rocket engines in use today are chemical rockets; that is, they obtain the energy needed to generate thrust by combining reactive chemicals to create a hot gas that is expanded to produce thrust. A significant limitation of chemical propulsion is that it has a relatively low specific impulse. Numerous concepts for advanced propulsion technologies with significantly higher values of specific impulse have been developed over the past 50 years. Advanced in-space propulsion technologies will enable much more effective exploration of our solar system, near and far, and will permit mission designers to plan missions to "fly anytime, anywhere, and complete a host of science objectives at the destinations" with greater reliability and safety. With a wide range of possible missions and candidate propulsion technologies with very diverse characteristics, the question of which technologies are 'best' for future missions is a difficult one. A portfolio of technologies to allow optimum propulsion solutions for a diverse set of missions and destinations are described in the roadmap and herein.

Induction Heating Model of Cermet Fuel Element Environmental Test (CFEET)

NASA Technical Reports Server (NTRS)

Gomez, Carlos F.; Bradley, D. E.; Cavender, D. P.; Mireles, O. R.; Hickman, R. R.; Trent, D.; Stewart, E.

2013-01-01

Deep space missions with large payloads require high specific impulse and relatively high thrust to achieve mission goals in reasonable time frames. Nuclear Thermal Rockets (NTR) are capable of producing a high specific impulse by employing heat produced by a fission reactor to heat and therefore accelerate hydrogen through a rocket nozzle providing thrust. Fuel element temperatures are very high (up to 3000 K) and hydrogen is highly reactive with most materials at high temperatures. Data covering the effects of high-temperature hydrogen exposure on fuel elements are limited. The primary concern is the mechanical failure of fuel elements due to large thermal gradients; therefore, high-melting-point ceramics-metallic matrix composites (cermets) are one of the fuels under consideration as part of the Nuclear Cryogenic Propulsion Stage (NCPS) Advance Exploration System (AES) technology project at the Marshall Space Flight Center. The purpose of testing and analytical modeling is to determine their ability to survive and maintain thermal performance in a prototypical NTR reactor environment of exposure to hydrogen at very high temperatures and obtain data to assess the properties of the non-nuclear support materials. The fission process and the resulting heating performance are well known and do not require that active fissile material to be integrated in this testing. A small-scale test bed; Compact Fuel Element Environmental Tester (CFEET), designed to heat fuel element samples via induction heating and expose samples to hydrogen is being developed at MSFC to assist in optimal material and manufacturing process selection without utilizing fissile material. This paper details the analytical approach to help design and optimize the test bed using COMSOL Multiphysics for predicting thermal gradients induced by electromagnetic heating (Induction heating) and Thermal Desktop for radiation calculations.

Strategies for the sustained human exploration of Mars

NASA Astrophysics Data System (ADS)

Landau, Damon Frederick

A variety of mission scenarios are compared in this thesis to assess the strengths and weaknesses of options for Mars exploration. The mission design space is modeled along two dimensions: trajectory architectures and propulsion system technologies. Direct, semi-direct, stop-over, semi-cycler, and cycler architectures are examined, and electric propulsion, nuclear thermal rockets, methane and oxygen production on Mars, Mars water excavation, aerocapture, and reusable propulsion systems are included in the technology assessment. The mission sensitivity to crew size, vehicle masses, and crew travel time is also examined. The primary figure of merit for a mission scenario is the injected mass to low-Earth orbit (IMLEO), though technology readiness levels (TRL) are also included. Several elements in the architecture dimension are explored in more detail. The Earth-Mars semi-cycler architecture is introduced and five families of Earth-Mars semi-cycler trajectories are presented along with optimized itineraries. Optimized cycler trajectories are also presented. In addition to Earth-Mars semi-cycler and cycler trajectories, conjunction-class, free-return, Mars-Earth semi-cycler, and low-thrust trajectories are calculated. Design parameters for optimal DeltaV trajectories are provided over a range of flight times (from 120 to 270 days) and launch years (between 2009 and 2022). Unlike impulsive transfers, the mass-optimal low-thrust trajectory depends strongly on the thrust and specific impulse of the propulsion system. A low-thrust version of the rocket equation is provided where the initial mass or thrust may be minimized by varying the initial acceleration and specific impulse. Planet-centered operations are also examined. A method to rotate a parking orbit about the line of apsides to achieve the proper orientation at departure is discussed, thus coupling the effects of parking-orbit orientation with the interplanetary trajectories. Also, a guidance algorithm for rendezvous during flybys in semi-cycler and cycler missions is presented with a control law for final approach. A forty-year plan to establish a permanent base on Mars is detailed and methods to expand the base are discussed. Once a large base is established, one-, two-, or three-vehicle systems may sustain the colonization of Mars.

Closed-loop thrust and pressure profile throttling of a nitrous oxide/hydroxyl-terminated polybutadiene hybrid rocket motor

NASA Astrophysics Data System (ADS)

Peterson, Zachary W.

Hybrid motors that employ non-toxic, non-explosive components with a liquid oxidizer and a solid hydrocarbon fuel grain have inherently safe operating characteristics. The inherent safety of hybrid rocket motors offers the potential to greatly reduce overall operating costs. Another key advantage of hybrid rocket motors is the potential for in-flight shutdown, restart, and throttle by controlling the pressure drop between the oxidizer tank and the injector. This research designed, developed, and ground tested a closed-loop throttle controller for a hybrid rocket motor using nitrous oxide and hydroxyl-terminated polybutadiene as propellants. The research simultaneously developed closed-loop throttle algorithms and lab scale motor hardware to evaluate the fidelity of the throttle simulations and algorithms. Initial open-loop motor tests were performed to better classify system parameters and to validate motor performance values. Deep-throttle open-loop tests evaluated limits of stable thrust that can be achieved on the test hardware. Open-loop tests demonstrated the ability to throttle the motor to less than 10% of maximum thrust with little reduction in effective specific impulse and acoustical stability. Following the open-loop development, closed-loop, hardware-in-the-loop tests were performed. The closed-loop controller successfully tracked prescribed step and ramp command profiles with a high degree of fidelity. Steady-state accuracy was greatly improved over uncontrolled thrust.

Experimental investigation of solid rocket motors for small sounding rockets

NASA Astrophysics Data System (ADS)

Suksila, Thada

2018-01-01

Experimentation and research of solid rocket motors are important subjects for aerospace engineering students. However, many institutes in Thailand rarely include experiments on solid rocket motors in research projects of aerospace engineering students, mainly because of the complexity of mixing the explosive propellants. This paper focuses on the design and construction of a solid rocket motor for total impulse in the class I-J that can be utilised as a small sounding rocket by researchers in the near future. Initially, the test stands intended for measuring the pressure in the combustion chamber and the thrust of the solid rocket motor were designed and constructed. The basic design of the propellant configuration was evaluated. Several formulas and ratios of solid propellants were compared for achieving the maximum thrust. The convenience of manufacturing and casting of the fabricated solid rocket motors were a critical consideration. The motor structural analysis such as the combustion chamber wall thickness was also discussed. Several types of nozzles were compared and evaluated for ensuring the maximum thrust of the solid rocket motors during the experiments. The theory of heat transfer analysis in the combustion chamber was discussed and compared with the experimental data.

Nuclear design of a vapor core reactor for space nuclear propulsion

NASA Astrophysics Data System (ADS)

Dugan, Edward T.; Watanabe, Yoichi; Kuras, Stephen A.; Maya, Isaac; Diaz, Nils J.

1993-01-01

Neutronic analysis methodology and results are presented for the nuclear design of a vapor core reactor for space nuclear propulsion. The Nuclear Vapor Thermal Reactor (NVTR) Rocket Engine uses modified NERVA geometry and systems which the solid fuel replaced by uranium tetrafluoride vapor. The NVTR is an intermediate term gas core thermal rocket engine with specific impulse in the range of 1000-1200 seconds; a thrust of 75,000 lbs for a hydrogen flow rate of 30 kg/s; average core exit temperatures of 3100 K to 3400 K; and reactor thermal powers of 1400 to 1800 MW. Initial calculations were performed on epithermal NVTRs using ZrC fuel elements. Studies are now directed at thermal NVTRs that use fuel elements made of C-C composite. The large ZrC-moderated reactors resulted in thrust-to-weight ratios of only 1 to 2; the compact C-C composite systems yield thrust-to-weight ratios of 3 to 5.

Nuclear Thermal Rocket (NTR) Propulsion and Power Systems for Outer Planetary Exploration Missions

NASA Technical Reports Server (NTRS)

Borowski, S. K.; Cataldo, R. L.

2001-01-01

The high specific impulse (I (sub sp)) and engine thrust generated using liquid hydrogen (LH2)-cooled Nuclear Thermal Rocket (NTR) propulsion makes them attractive for upper stage applications for difficult robotic science missions to the outer planets. Besides high (I (sub sp)) and thrust, NTR engines can also be designed for "bimodal" operation allowing substantial amounts of electrical power (10's of kWe ) to be generated for onboard spacecraft systems and high data rate communications with Earth during the course of the mission. Two possible options for using the NTR are examined here. A high performance injection stage utilizing a single 15 klbf thrust engine can inject large payloads to the outer planets using a 20 t-class launch vehicle when operated in an "expendable mode". A smaller bimodal NTR stage generating approx. 1 klbf of thrust and 20 to 40 kWe for electric propulsion can deliver approx. 100 kg using lower cost launch vehicles. Additional information is contained in the original extended abstract.

Comparison of two procedures for predicting rocket engine nozzle performance

NASA Technical Reports Server (NTRS)

Davidian, Kenneth J.

1987-01-01

Two nozzle performance prediction procedures which are based on the standardized JANNAF methodology are presented and compared for four rocket engine nozzles. The first procedure required operator intercedence to transfer data between the individual performance programs. The second procedure is more automated in that all necessary programs are collected into a single computer code, thereby eliminating the need for data reformatting. Results from both procedures show similar trends but quantitative differences. Agreement was best in the predictions of specific impulse and local skin friction coefficient. Other compared quantities include characteristic velocity, thrust coefficient, thrust decrement, boundary layer displacement thickness, momentum thickness, and heat loss rate to the wall. Effects of wall temperature profile used as an input to the programs was investigated by running three wall temperature profiles. It was found that this change greatly affected the boundary layer displacement thickness and heat loss to the wall. The other quantities, however, were not drastically affected by the wall temperature profile change.

Propulsion requirements for reusable single-stage-to-orbit rocket vehicles

NASA Astrophysics Data System (ADS)

Stanley, Douglas O.; Engelund, Walter C.; Lepsch, Roger

1994-05-01

The conceptual design of a single-stage-to-orbit (SSTO) vehicle using a wide variety of evolutionary technologies has recently been completed as a part of NASA's Advanced Manned Launch System (AMLS) study. The employment of new propulsion system technologies is critical to the design of a reasonably sized, operationally efficient SSTO vehicle. This paper presents the propulsion system requirements identified for this near-term AMLS SSTO vehicle. Sensitivities of the vehicle to changes in specific impulse and sea-level thrust-to-weight ratio are examined. The results of a variety of vehicle/propulsion system trades performed on the near-term AMLS SSTO vehicle are also presented.

Space Shuttle propulsion performance reconstruction from flight data

NASA Technical Reports Server (NTRS)

Rogers, Robert M.

1989-01-01

The aplication of extended Kalman filtering to estimating Space Shuttle Solid Rocket Booster (SRB) performance, specific impulse, from flight data in a post-flight processing computer program. The flight data used includes inertial platform acceleration, SRB head pressure, and ground based radar tracking data. The key feature in this application is the model used for the SRBs, which represents a reference quasi-static internal ballistics model normalized to the propellant burn depth. Dynamic states of mass overboard and propellant burn depth are included in the filter model to account for real-time deviations from the reference model used. Aerodynamic, plume, wind and main engine uncertainties are included.

Accuracy Improvement in Magnetic Field Modeling for an Axisymmetric Electromagnet

NASA Technical Reports Server (NTRS)

Ilin, Andrew V.; Chang-Diaz, Franklin R.; Gurieva, Yana L.; Il,in, Valery P.

2000-01-01

This paper examines the accuracy and calculation speed for the magnetic field computation in an axisymmetric electromagnet. Different numerical techniques, based on an adaptive nonuniform grid, high order finite difference approximations, and semi-analitical calculation of boundary conditions are considered. These techniques are being applied to the modeling of the Variable Specific Impulse Magnetoplasma Rocket. For high-accuracy calculations, a fourth-order scheme offers dramatic advantages over a second order scheme. For complex physical configurations of interest in plasma propulsion, a second-order scheme with nonuniform mesh gives the best results. Also, the relative advantages of various methods are described when the speed of computation is an important consideration.

Metallic Hydrogen - Potentially a High Energy Rocket Propellant

NASA Technical Reports Server (NTRS)

Cole, John; Silvera, Ike

2007-01-01

Pure metallic hydrogen is predicted to have a specific impulse (Isp) of 1700 seconds, but the reaction temperature is too high for current engine materials. Diluting metallic hydrogen with liquid hydrogen can reduce the reaction temperature to levels compatible with current material limits and still provide an Isp greater than 900 s. Metallic hydrogen has not yet been produced on earth, but experimental techniques exist that may change this situation. This paper will provide a brief description of metallic hydrogen and the status of experiments that may soon produce detectable quantities of this material in the lab. Also provided are some characteristics for diluted metallic hydrogen engines and launch vehicles.

The microspace launcher: first step to the fully air-breathing space launcher

NASA Astrophysics Data System (ADS)

Falempin, F.; Bouchez, M.; Calabro, M.

2009-09-01

A possible application for the high-speed air-breathing propulsion is the fully or partially reusable space launcher. Indeed, by combining the high-speed air-breathing propulsion with a conventional rocket engine (combined cycle or combined propulsion system), it should be possible to improve the average installed specific impulse along the ascent trajectory and then make possible more performing launchers and, hopefully, a fully reusable one. During the last 15 years, a lot of system studies have been performed in France on that subject within the framework of different and consecutive programs. Nevertheless, these studies never clearly demonstrated that a space launcher could take advantage of using a combined propulsion system. During last years, the interest to air-breathing propulsion for space application has been revisited. During this review and taking into account technologies development activities already in progress in Europe, clear priorities have been identified regarding a minimum complementary research and technology program addressing specific needs of space launcher application. It was also clearly identified that there is the need to restart system studies taking advantage of recent progress made regarding knowledge, tools, and technology and focusing on more innovative airframe/propulsion system concepts enabling better trade-off between structural efficiency and propulsion system performance. In that field, a fully axisymmetric configuration has been considered for a microspace launcher (10 kg payload). The vehicle is based on a main stage powered by air-breathing propulsion, combined or not with liquid rocket mode. A "kick stage," powered by a solid rocket engine provides the final acceleration. A preliminary design has been performed for different variants: one using a separated booster and a purely air-breathing main stage, a second one using a booster and a main stage combining air-breathing and rocket mode, a third one without separated booster, the main stage ensuring the initial acceleration in liquid rocket mode and a complementary acceleration phase in rocket mode beyond the air-breathing propulsion system operation. Finally, the liquid rocket engine of this third variant can be replaced by a continuous detonation wave rocket engine. The paper describes the main guidelines for the design of these variants and provides their main characteristics. On this basis, the achievable performance, estimated by trajectory simulation, are detailed.

Small Launch Vehicle Concept Development for Affordable Multi-Stage Inline Configurations

NASA Technical Reports Server (NTRS)

Beers, Benjamin R.; Waters, Eric D.; Philips, Alan D.; Threet, Grady E., Jr.

2014-01-01

The Advanced Concepts Office at NASA's George C. Marshall Space Flight Center conducted a study of two configurations of a three-stage, inline, liquid propellant small launch vehicle concept developed on the premise of maximizing affordability by targeting a specific payload capability range based on current and future industry demand. The initial configuration, NESC-1, employed liquid oxygen as the oxidizer and rocket propellant grade kerosene as the fuel in all three stages. The second and more heavily studied configuration, NESC-4, employed liquid oxygen and rocket propellant grade kerosene on the first and second stages and liquid oxygen and liquid methane fuel on the third stage. On both vehicles, sensitivity studies were first conducted on specific impulse and stage propellant mass fraction in order to baseline gear ratios and drive the focus of concept development. Subsequent sensitivity and trade studies on the NESC-4 concept investigated potential impacts to affordability due to changes in gross liftoff mass and/or vehicle complexity. Results are discussed at a high level to understand the impact severity of certain sensitivities and how those trade studies conducted can either affect cost, performance, or both.

Partial admission effect on the performance and vibration of a supersonic impulse turbine

NASA Astrophysics Data System (ADS)

Lee, Hang Gi; Shin, Ju Hyun; Choi, Chang-Ho; Jeong, Eunhwan; Kwon, Sejin

2018-04-01

This study experimentally investigates the effects of partial admission on the performance and vibration outcomes of a supersonic impulse turbine with circular nozzles. The turbine of a turbopump for a gas-generator-type liquid rocket engine in the Korea Space Launch Vehicle-II is of the supersonic impulse type with the partial admission configuration for obtaining a high specific power. Partial admission turbines with a low-flow-rate working gas exhibit benefits over turbines with full admission, such as loss reduction, ease of controllability of the turbine power output, and simple turbine configurations with separate starting sections. However, the radial force of the turbine rotor due to the partial admission causes an increase in turbine vibration. Few experimental studies have previously been conducted regarding the partial admission effects on supersonic impulse turbines with circular nozzles. In the present study, performance tests of supersonic impulse turbines with circular nozzles were conducted for various partial admission ratios using a turbine test facility with high-pressure air in order to investigate the resulting aerodynamic performance and vibration. Four types of turbines with partial admission ratios of 0.17, 0.42, 0.75 and 0.83 were tested. Results show that the efficiencies at the design point increase linearly as the partial admission ratios increase. Moreover, as the velocity ratios increase, the difference in efficiency from the reference turbine with a partial admission ratio of 0.83 becomes increasingly significant, and the magnitudes of these differences are proportional to the square of the velocity ratios. Likewise, the decrease in the partial admission ratio results in an increase in the turbine vibration level owing to the increase in the radial force.

Fortran 4 program for two-impulse rendezvous analysis

NASA Technical Reports Server (NTRS)

Barling, W. H., Jr.; Brothers, W. J.; Darling, W. H., Jr.

1967-01-01

Program determines if rendezvous in near space is possible, and performs an analysis to determine the approximate required values of the magnitude and direction of two thrust applications of the upper stage of a rocket firing. The analysis is performed by using ordinary Keplerian mechanics.

Ascent performance feasibility for next-generation spacecraft

NASA Astrophysics Data System (ADS)

Mancuso, Salvatore Massimo

This thesis deals with the optimization of the ascent trajectories for single-stage suborbital (SSSO), single-stage-to-orbit (SSTO), and two-stage-to-orbit (TSTO) rocket-powered spacecraft. The maximum payload weight problem has been solved using the sequential gradient-restoration algorithm. For the TSTO case, some modifications to the original version of the algorithm have been necessary in order to deal with discontinuities due to staging and the fact that the functional being minimized depends on interface conditions. The optimization problem is studied for different values of the initial thrust-to-weight ratio in the range 1.3 to 1.6, engine specific impulse in the range 400 to 500 sec, and spacecraft structural factor in the range 0.08 to 0.12. For the TSTO configuration, two subproblems are studied: uniform structural factor between stages and nonuniform structural factor between stages. Due to the regular behavior of the results obtained, engineering approximations have been developed which connect the maximum payload weight to the engine specific impulse and spacecraft structural factor; in turn, this leads to useful design considerations. Also, performance sensitivity to the scale of the aerodynamic drag is studied, and it is shown that its effect on payload weight is relatively small, even for drag changes approaching ± 50%. The main conclusions are that: the design of a SSSO configuration appears to be feasible; the design of a SSTO configuration might be comfortably feasible, marginally feasible, or unfeasible, depending on the parameter values assumed; the design of a TSTO configuration is not only feasible, but its payload appears to be considerably larger than that of a SSTO configuration. Improvements in engine specific impulse and spacecraft structural factor are desirable and crucial for SSTO feasibility; indeed, it appears that aerodynamic improvements do not yield significant improvements in payload weight.

Effect of the Thruster Configurations on a Laser Ignition Microthruster

NASA Astrophysics Data System (ADS)

Koizumi, Hiroyuki; Hamasaki, Kyoichi; Kondo, Ryo; Okada, Keisuke; Nakano, Masakatsu; Arakawa, Yoshihiro

Research and development of small spacecraft have advanced extensively throughout the world and propulsion devices suitable for the small spacecraft, microthruster, is eagerly anticipated. The authors proposed a microthruster using 1—10-mm-size solid propellant. Small pellets of solid propellant are installed in small combustion chambers and ignited by the irradiation of diode laser beam. This thruster is referred as to a laser ignition microthruster. Solid propellant enables large thrust capability and compact propulsion system. To date theories of a solid-propellant rocket have been well established. However, those theories are for a large-size solid propellant and there are a few theories and experiments for a micro-solid rocket of 1—10mm class. This causes the difficulty of the optimum design of a micro-solid rocket. In this study, we have experimentally investigated the effect of thruster configurations on a laser ignition microthruster. The examined parameters are aperture ratio of the nozzle, length of the combustion chamber, area of the nozzle throat, and divergence angle of the nozzle. Specific impulse dependences on those parameters were evaluated. It was found that large fraction of the uncombusted propellant was the main cause of the degrading performance. Decreasing the orifice diameter in the nozzle with a constant open aperture ratio was an effective method to improve this degradation.

Predicting performance of axial pump inducer of LOX booster turbo-pump of staged combustion cycle based rocket engine using CFD

NASA Astrophysics Data System (ADS)

Mishra, Arpit; Ghosh, Parthasarathi

2015-12-01

For low cost, high thrust, space missions with high specific impulse and high reliability, inert weight needs to be minimized and thereby increasing the delivered payload. Turbopump feed system for a liquid propellant rocket engine (LPRE) has the highest power to weight ratio. Turbopumps are primarily equipped with an axial flow inducer to achieve the high angular velocity and low suction pressure in combination with increased system reliability. Performance of the turbopump strongly depends on the performance of the inducer. Thus, for designing a LPRE turbopump, demands optimization of the inducer geometry based on the performance of different off-design operating regimes. In this paper, steady-state CFD analysis of the inducer of a liquid oxygen (LOX) axial pump used as a booster pump for an oxygen rich staged combustion cycle rocket engine has been presented using ANSYS® CFX. Attempts have been made to obtain the performance characteristic curves for the LOX pump inducer. The formalism has been used to predict the performance of the inducer for the throttling range varying from 80% to 113% of nominal thrust and for the different rotational velocities from 4500 to 7500 rpm. The results have been analysed to determine the region of cavitation inception for different inlet pressure.

Low Cost Nuclear Thermal Rocket Cermet Fuel Element Environment Testing

NASA Technical Reports Server (NTRS)

Bradley, David E.; Mireles, Omar R.; Hickman, Robert R.

2011-01-01

Deep space missions with large payloads require high specific impulse (Isp) and relatively high thrust in order to achieve mission goals in reasonable time frames. Conventional, storable propellants produce average Isp. Nuclear thermal rockets (NTR) capable of high Isp thrust have been proposed. NTR employs heat produced by fission reaction to heat and therefore accelerate hydrogen which is then forced through a rocket nozzle providing thrust. Fuel element temperatures are very high (up to 3000K) and hydrogen is highly reactive with most materials at high temperatures. Data covering the effects of high temperature hydrogen exposure on fuel elements is limited. The primary concern is the mechanical failure of fuel elements which employ high-melting-point metals, ceramics or a combination (cermet) as a structural matrix into which the nuclear fuel is distributed. It is not necessary to include fissile material in test samples intended to explore high temperature hydrogen exposure of the structural support matrices. A small-scale test bed designed to heat fuel element samples via non-contact RF heating and expose samples to hydrogen is being developed to assist in optimal material and manufacturing process selection without employing fissile material. This paper details the test bed design and results of testing conducted to date.

Studies on an aerial propellant transfer space plane (APTSP)

NASA Astrophysics Data System (ADS)

Jayan, N.; Biju Kumar, K. S.; Gupta, Anish Kumar; Kashyap, Akhilesh Kumar; Venkatraman, Kartik; Mathew, Joseph; Mukunda, H. S.

2004-04-01

This paper presents a study of a fully reusable earth-to-orbit launch vehicle concept with horizontal take-off and landing, employing a turbojet engine for low speed, and a rocket for high-speed acceleration and space operations. This concept uses existing technology to the maximum possible extent, thereby reducing development time, cost and effort. It uses the experience in aerial filling of military aircrafts for propellant filling at an altitude of 13 km at a flight speed of M=0.85. Aerial filling of propellant reduces the take-off weight significantly thereby minimizing the structural weight of the vehicle. The vehicle takes off horizontally and uses turbojet engines till the end of the propellant filling operation. The rocket engines provide thrust for the next phase till the injection of a satellite at LEO. A sensitivity analysis of the mission with respect to rocket engine specific impulse and overall vehicle structural factor is also presented in this paper. A conceptual design of space plane with a payload capability of 10 ton to LEO is carried out. The study shows that the realization of an aerial propellant transfer space plane is possible with limited development of new technology thus reducing the demands on the finances required for achieving the objectives.

Design Considerations for Clean QED Fusion Propulsion Systems

NASA Astrophysics Data System (ADS)

Bussard, Robert W.; Jameson, Lorin W.

1994-07-01

The direct production of electric power appears possible from fusion reactions between fuels whose products consist solely of charged particles and thus do not present radiation hazards from energetic neutron production, as do reactions involving deuteron-bearing fuels. Among these are the fuels p, 11B, 3He, and 6Li. All of these can be ``burned'' in inertial-electrostatic-fusion (IEF) devices to power QED fusion-electric rocket engines. These IEF sources provide direct-converted electrical power at high voltage (MeV) to drive e-beams for efficient propellant heating to extreme temperatures, with resulting high specific impulse performance capabilities. IEF/QED engine systems using p11B can outperform all other advanced concepts for controlled fusion propulsion by 2-3 orders of magnitude, while 6Li6Li fusion yields one order of magnitude less advance. Either of these fusion rocket propulsion systems can provide very rapid transit for solar system missions, with high payload fractions in single-stage vehicles. The 3He3He reaction can not be used practically for direct electric conversion because of the wide spread in energy of its fusion products. However, it may eventually prove useful for thermal/electrical power generation in central station power plants, or for direct-fusion-product (DFP) propellant heatingin advanced deep-space rocket engines.

A Nuclear Cryogenic Propulsion Stage for Near-Term Space Missions

NASA Technical Reports Server (NTRS)

Houts, Michael G.; Kim, Tony; Emrich, William J.; Hickman, Robert R.; Broadway, Jeramie W.; Gerrish, Harold P.; Adams, Robert B.; Bechtel, Ryan D.; Borowski, Stanley K.; George, Jeffrey A.

2013-01-01

The potential capability of NTP is game changing for space exploration. A first generation NCPS could provide high thrust at a specific impulse above 900 s, roughly double that of state of the art chemical engines. Near-term NCPS systems would provide a foundation for the development of significantly more advanced, higher performance systems. John F. Kennedy made his historic special address to Congress on the importance of space on May 25, 1961, "First, I believe that this nation should commit itself to achieving the goal, before this decade is out, of landing a man on the Moon and returning him safely to the Earth..." This was accomplished. John F. Kennedy also made a second request, "Secondly... accelerate development of the Rover nuclear rocket. This gives promise of some day providing a means for even more exciting and ambitious exploration of space, perhaps beyond the Moon, perhaps to the very end of the solar system itself." The investment in the Rover nuclear rocket program provided the foundation of technology that gives us assurance for greater performing rockets that are capable of taking us further into space. Combined with current technologies, the vision to go beyond the Moon and to the very end of the solar system can be realized with space nuclear propulsion and power.

Numerical simulations on unsteady operation processes of N2O/HTPB hybrid rocket motor with/without diaphragm

NASA Astrophysics Data System (ADS)

Zhang, Shuai; Hu, Fan; Wang, Donghui; Okolo. N, Patrick; Zhang, Weihua

2017-07-01

Numerical simulations on processes within a hybrid rocket motor were conducted in the past, where most of these simulations carried out majorly focused on steady state analysis. Solid fuel regression rate strongly depends on complicated physicochemical processes and internal fluid dynamic behavior within the rocket motor, which changes with both space and time during its operation, and are therefore more unsteady in characteristics. Numerical simulations on the unsteady operational processes of N2O/HTPB hybrid rocket motor with and without diaphragm are conducted within this research paper. A numerical model is established based on two dimensional axisymmetric unsteady Navier-Stokes equations having turbulence, combustion and coupled gas/solid phase formulations. Discrete phase model is used to simulate injection and vaporization of the liquid oxidizer. A dynamic mesh technique is applied to the non-uniform regression of fuel grain, while results of unsteady flow field, variation of regression rate distribution with time, regression process of burning surface and internal ballistics are all obtained. Due to presence of eddy flow, the diaphragm increases regression rate further downstream. Peak regression rates are observed close to flow reattachment regions, while these peak values decrease gradually, and peak position shift further downstream with time advancement. Motor performance is analyzed accordingly, and it is noticed that the case with diaphragm included results in combustion efficiency and specific impulse efficiency increase of roughly 10%, and ground thrust increase of 17.8%.

Design of a 500 lbf liquid oxygen and liquid methane rocket engine for suborbital flight

NASA Astrophysics Data System (ADS)

Trillo, Jesus Eduardo

Liquid methane (LCH4)is the most promising rocket fuel for our journey to Mars and other space entities. Compared to liquid hydrogen, the most common cryogenic fuel used today, methane is denser and can be stored at a more manageable temperature; leading to more affordable tanks and a lighter system. The most important advantage is it can be produced from local sources using in-situ resource utilization (ISRU) technology. This will allow the production of the fuel needed to come back to earth on the surface of Mars, or the space entity being explored, making the overall mission more cost effective by enabling larger usable mass. The major disadvantage methane has over hydrogen is it provides a lower specific impulse, or lower rocket performance. The UTEP Center for Space Exploration and Technology Research (cSETR) in partnership with the National Aeronautics and Space Administration (NASA) has been the leading research center for the advancement of Liquid Oxygen (LOX) and Liquid Methane (LCH4) propulsion technologies. Through this partnership, the CROME engine, a throattable 500 lbf LOX/LCH4 rocket engine, was designed and developed. The engine will serve as the main propulsion system for Daedalus, a suborbital demonstration vehicle being developed by the cSETR. The purpose of Daedalus mission and the engine is to fire in space under microgravity conditions to demonstrate its restartability. This thesis details the design process, decisions, and characteristics of the engine to serve as a complete design guide.

H.F. emission related to the Li+ ion beam injected into ionosphere - ``PLAZMA'' rocket experiment

NASA Astrophysics Data System (ADS)

Klos, Z.; Zbyszynski, Z.; Agafonov, U. F.; Managadze, G. G.; Mayorov, A. D.

1993-10-01

The H.F. emission generated by artificial ion beam injected into ionosphere was observed either with a wave detector and ion gun attached to the rocket through out the flight, or when the gun was deployed on subpayload. Generally the observations show unstructured shape of the H.F. spectrum. In the PLAZMA active rocket experiment - when ionospheric plasma was perturbed by the operation of impulse ion gun, which injected 300 A, 8.3 eV Li+ions - the waves in the 0.1 - 10 MHz frequency range were observed. The results have shown, that when the wave detector and the ion gun are attached to the rocket the emission enhances in the lower as well as in the upper parts of the spectrum. On the other hand only the lower increase is maintained when ion gun is removing away on the subpayload. The observed sequence of H.F. spectra is presented.

Studies of Fission Fragment Rocket Engine Propelled Spacecraft

NASA Technical Reports Server (NTRS)

Werka, Robert O.; Clark, Rodney; Sheldon, Rob; Percy, Thomas K.

2014-01-01

The NASA Office of Chief Technologist has funded from FY11 through FY14 successive studies of the physics, design, and spacecraft integration of a Fission Fragment Rocket Engine (FFRE) that directly converts the momentum of fission fragments continuously into spacecraft momentum at a theoretical specific impulse above one million seconds. While others have promised future propulsion advances if only you have the patience, the FFRE requires no waiting, no advances in physics and no advances in manufacturing processes. Such an engine unequivocally can create a new era of space exploration that can change spacecraft operation. The NIAC (NASA Institute for Advanced Concepts) Program Phase 1 study of FY11 first investigated how the revolutionary FFRE technology could be integrated into an advanced spacecraft. The FFRE combines existent technologies of low density fissioning dust trapped electrostatically and high field strength superconducting magnets for beam management. By organizing the nuclear core material to permit sufficient mean free path for escape of the fission fragments and by collimating the beam, this study showed the FFRE could convert nuclear power to thrust directly and efficiently at a delivered specific impulse of 527,000 seconds. The FY13 study showed that, without increasing the reactor power, adding a neutral gas to the fission fragment beam significantly increased the FFRE thrust through in a manner analogous to a jet engine afterburner. This frictional interaction of gas and beam resulted in an engine that continuously produced 1000 pound force of thrust at a delivered impulse of 32,000 seconds, thereby reducing the currently studied DRM 5 round trip mission to Mars from 3 years to 260 days. By decreasing the gas addition, this same engine can be tailored for much lower thrust at much higher impulse to match missions to more distant destinations. These studies created host spacecraft concepts configured for manned round trip journeys. While the vehicles are very large, they are primarily made up of a habitat payload on one end, the engine on the opposite end and a connecting spine containing radiator acreage needed to reject the heat of this powerful, but inefficient engine. These studies concluded that the engine and spacecraft are within today's technology, could be built, tested, launched on several SLS launchers, integrated, checked out, maintained at an in-space LEO base, and operated for decades just as Caribbean cruise ships operate today. The nuclear issues were found to be far less daunting that [than for] current nuclear engines. The FFRE produces very small amounts of radioactive efflux compared to their impulse, easily contained in an evacuated "bore-hole" test site. The engine poses no launch risk since it is simply a structure containing no fissionable material. The nuclear fuel is carried to orbit in containers highly crash-proofed for launch accidents from which it, in a liquid medium, is injected into the FFRE. The radioactive exhaust, with a velocity above 300 kilometers per second rapidly leaves the solar system.

High temperature thruster technology for spacecraft propulsion

NASA Technical Reports Server (NTRS)

Schneider, Steven J.

1991-01-01

A technology program intended to develop high-temperature oxidation-resistant thrusters for spacecraft applications is considered. The program will provide the requisite material characterizations and fabrication to incorporate iridium coated rhenium material into small rockets for spacecraft propulsion. This material increases the operating temperature of thrusters to 2200 C, a significant increase over the 1400 C of the silicide-coated niobium chambers currently used. Stationkeeping class 22 N engines fabricated from iridium-coated rhenium have demonstrated steady state specific impulses 20-25 seconds higher than niobium chambers. These improved performances are obtained by reducing or eliminating the fuel film cooling requirements in the combustion chamber while operating at the same overall mixture ratio as conventional engines.

Performance potential of air turbo-ramjet employing supersonic through-flow fan

NASA Technical Reports Server (NTRS)

Kepler, C. E.; Champagne, G. A.

1989-01-01

A study was conducted to assess the performance potential of a supersonic through-flow fan in an advanced engine designed to power a Mach-5 cruise vehicle. It included a preliminary evaluation of fan performance requirements and the desirability of supersonic versus subsonic combustion, the design and performance of supersonic fans, and the conceptual design of a single-pass air-turbo-rocket/ramjet engine for a Mach 5 cruise vehicle. The study results showed that such an engine could provide high thrust over the entire speed range from sea-level takeoff to Mach 5 cruise, especially over the transonic speed range, and high fuel specific impulse at the Mach 5 cruise condition, with the fan windmilling.

Solar thermal propulsion for planetary spacecraft

NASA Technical Reports Server (NTRS)

Sercel, J. C.

1985-01-01

Previous studies have shown that many desirable planetary exploration missions require large injection delta-V. Solar Thermal Rocket (STR) propulsion, under study for orbit-raising applications may enhance or enable such high-energy missions. The required technology of thermal control for liquid hydrogen propellant is available for the required storage duration. Self-deploying, inflatable solar concentrators are under study. The mass penalty for passive cryogenic thermal control, liquid hydrogen tanks and solar concentrators does not compromise the specific impulse advantage afforded by the STR as compared to chemical propulsion systems. An STR injection module is characterized and performance is evaluated by comparison to electric propulsion options for the Saturn Orbiter Titan Probe (SOTP) and Uranus Flyby Uranus Probe (UFUP) missions.

Characterization of the space shuttle reaction control system engine

NASA Technical Reports Server (NTRS)

Wilson, M. S.; Stechman, R. C.; Edelman, R. B.; Fortune, O. F.; Economos, C.

1972-01-01

A computer program was developed and written in FORTRAN 5 which predicts the transient and steady state performance and heat transfer characteristics of a pulsing GO2/GH2 rocket engine. This program predicts the dynamic flow and ignition characteristics which, when combined in a quasi-steady state manner with the combustion and mixing analysis program, will provide the thrust and specific impulse of the engine as a function of time. The program also predicts the transient and steady state heat transfer characteristics of the engine using various cooling concepts. The computer program, test case, and documentation are presented. The program is applicable to any system capable of utilizing the FORTRAN 4 or FORTRAN 5 language.

Gaseous fuel nuclear reactor research

NASA Technical Reports Server (NTRS)

Schwenk, F. C.; Thom, K.

1975-01-01

Gaseous-fuel nuclear reactors are described; their distinguishing feature is the use of fissile fuels in a gaseous or plasma state, thereby breaking the barrier of temperature imposed by solid-fuel elements. This property creates a reactor heat source that may be able to heat the propellant of a rocket engine to 10,000 or 20,000 K. At this temperature level, gas-core reactors would provide the breakthrough in propulsion needed to open the entire solar system to manned and unmanned spacecraft. The possibility of fuel recycling makes possible efficiencies of up to 65% and nuclear safety at reduced cost, as well as high-thrust propulsion capabilities with specific impulse up to 5000 sec.

Electrolysis Propulsion Provides High-Performance, Inexpensive, Clean Spacecraft Propulsion

NASA Technical Reports Server (NTRS)

deGroot, Wim A.

1999-01-01

An electrolysis propulsion system consumes electrical energy to decompose water into hydrogen and oxygen. These gases are stored in separate tanks and used when needed in gaseous bipropellant thrusters for spacecraft propulsion. The propellant and combustion products are clean and nontoxic. As a result, costs associated with testing, handling, and launching can be an order of magnitude lower than for conventional propulsion systems, making electrolysis a cost-effective alternative to state-of-the-art systems. The electrical conversion efficiency is high (>85 percent), and maximum thrust-to-power ratios of 0.2 newtons per kilowatt (N/kW), a 370-sec specific impulse, can be obtained. A further advantage of the water rocket is its dual-mode potential. For relatively high thrust applications, the system can be used as a bipropellant engine. For low thrust levels and/or small impulse bit requirements, cold gas oxygen can be used alone. An added innovation is that the same hardware, with modest modifications, can be converted into an energy-storage and power-generation fuel cell, reducing the spacecraft power and propulsion system weight by an order of magnitude.

SPOC Benchmark Case: SNRE Model

DOE Office of Scientific and Technical Information (OSTI.GOV)

Vishal Patel; Michael Eades; Claude Russel Joyner II

The Small Nuclear Rocket Engine (SNRE) was modeled in the Center for Space Nuclear Research’s (CSNR) Space Propulsion Optimization Code (SPOC). SPOC aims to create nuclear thermal propulsion (NTP) geometries quickly to perform parametric studies on design spaces of historic and new NTP designs. The SNRE geometry was modeled in SPOC and a critical core with a reasonable amount of criticality margin was found. The fuel, tie-tubes, reflector, and control drum masses were predicted rather well. These are all very important for neutronics calculations so the active reactor geometries created with SPOC can continue to be trusted. Thermal calculations ofmore » the average and hot fuel channels agreed very well. The specific impulse calculations used historically and in SPOC disagree so mass flow rates and impulses differed. Modeling peripheral and power balance components that do not affect nuclear characteristics of the core is not a feature of SPOC and as such, these components should continue to be designed using other tools. A full paper detailing the available SNRE data and comparisons with SPOC outputs will be submitted as a follow-up to this abstract.« less

Affordable Development and Demonstration of a Small Nuclear Thermal Rocket (NTR) Engine and Stage: How Small Is Big Enough?

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Sefcik, Robert J.; Fittje, James E.; McCurdy, David R.; Qualls, Arthur L.; Schnitzler, Bruce G.; Werner, James E.; Weitzberg, Abraham; Joyner, Claude R.

2016-01-01

The Nuclear Thermal Rocket (NTR) derives its energy from fission of uranium-235 atoms contained within fuel elements that comprise the engine's reactor core. It generates high thrust and has a specific impulse potential of approximately 900 specific impulse - a 100 percent increase over today's best chemical rockets. The Nuclear Thermal Propulsion (NTP) project, funded by NASA's Advanced Exploration Systems (AES) program, includes five key task activities: (1) Recapture, demonstration, and validation of heritage graphite composite (GC) fuel (selected as the Lead Fuel option); (2) Engine Conceptual Design; (3) Operating Requirements Definition; (4) Identification of Affordable Options for Ground Testing; and (5) Formulation of an Affordable Development Strategy. During fiscal year (FY) 2014, a preliminary Design Development Test and Evaluation (DDT&E) plan and schedule for NTP development was outlined by the NASA Glenn Research Center (GRC), Department of Energy (DOE) and industry that involved significant system-level demonstration projects that included Ground Technology Demonstration (GTD) tests at the Nevada National Security Site (NNSS), followed by a Flight Technology Demonstration (FTD) mission. To reduce cost for the GTD tests and FTD mission, small NTR engines, in either the 7.5 or 16.5 kilopound-force thrust class, were considered. Both engine options used GC fuel and a common fuel element (FE) design. The small approximately 7.5 kilopound-force criticality-limited engine produces approximately157 thermal megawatts and its core is configured with parallel rows of hexagonal-shaped FEs and tie tubes (TTs) with a FE to TT ratio of approximately 1:1. The larger approximately 16.5 kilopound-force Small Nuclear Rocket Engine (SNRE), developed by Los Alamos National Laboratory (LANL) at the end of the Rover program, produces approximately 367 thermal megawatts and has a FE to TT ratio of approximately 2:1. Although both engines use a common 35-inch (approximately 89-centimeters) -long FE, the SNRE's larger diameter core contains approximately 300 more FEs needed to produce an additional 210 thermal megawatts of power. To reduce the cost of the FTD mission, a simple one-burn lunar flyby mission was considered to reduce the liquid hydrogen (LH2) propellant loading, the stage size and complexity. Use of existing and flight proven liquid rocket and stage hardware (e.g., from the RL10B-2 engine and Delta Cryogenic Second Stage) was also maximized to further aid affordability. This paper examines the pros and cons of using these two small engine options, including their potential to support future human exploration missions to the Moon, near Earth asteroids (NEA), and Mars, and recommends a preferred size. It also provides a preliminary assessment of the key activities, development options, and schedule required to affordably build, ground test and fly a small NTR engine and stage within a 10-year timeframe.

Variable Specific Impulse Magnetoplasma Rocket Engine

NASA Technical Reports Server (NTRS)

Chang-Diaz, Franklin R. (Inventor)

2002-01-01

An engine is disclosed, including a controllable output plasma generator, a controllable heater for selectably raising a temperature of the plasma connected to an outlet of the plasma generator, and a nozzle connected to an outlet of the heater, through which heated plasma is discharged to provide thrust. In one embodiment, the source of plasma is a helicon generator. In one embodiment, the heater is an ion cyclotron resonator. In one embodiment, the nozzle is a radially diverging magnetic field disposed on a discharge side of the heater so that helically travelling particles in the beater exit the heater at high axial velocity. A particular embodiment includes control circuits for selectably directing a portion of radio frequency power from an RF generator to the helicon generator and to the cyclotron resonator so that the thrust output and the specific impulse of the engine can be selectively controlled. A method of propelling a vehicle is also disclosed. The method includes generating a plasma, heating said plasma, and discharging the heated plasma through a nozzle. In one embodiment, the nozzle is a diverging magnetic field. In this embodiment, the heating is performed by applying a radio frequency electro magnetic field to the plasma at the ion cyclotron frequency in an axially polarized DC magnetic field.

Investigation of Recombination Processes In A Magnetized Plasma

NASA Technical Reports Server (NTRS)

Chavers, Greg; Chang-Diaz, Franklin; Rodgers, Stephen L. (Technical Monitor)

2002-01-01

Interplanetary travel requires propulsion systems that can provide high specific impulse (Isp), while also having sufficient thrust to rapidly accelerate large payloads. One such propulsion system is the Variable Specific Impulse Magneto-plasma Rocket (VASIMR), which creates, heats, and exhausts plasma to provide variable thrust and Isp, optimally meeting the mission requirements. A large fraction of the energy to create the plasma is frozen in the exhaust in the form of ionization energy. This loss mechanism is common to all electromagnetic plasma thrusters and has an impact on their efficiency. When the device operates at high Isp, where the exhaust kinetic energy is high compared to the ionization energy, the frozen flow component is of little consequence; however, at low Isp, the effect of the frozen flow may be important. If some of this energy could be recovered through recombination processes, and re-injected as neutral kinetic energy, the efficiency of VASIMR, in its low Isp/high thrust mode may be improved. In this operating regime, the ionization energy is a large portion of the total plasma energy. An experiment is being conducted to investigate the possibility of recovering some of the energy used to create the plasma. This presentation will cover the progress and status of the experiment involving surface recombination of the plasma.

Nuclear-Rocket Propulsion

NASA Technical Reports Server (NTRS)

Rom, Frank E.

1968-01-01

The three basic types of nuclear power-plants (solid, liquid, and gas core) are compared on the bases of performance potential and the status of current technology. The solid-core systems are expected to have impulses in the range of 850 seconds, any thrust level (as long as it is greater than 10,000 pounds (44,480 newtons)), and thrust-to-engine-weight ratios of 2 to 20 pounds per pound (19.7 to 197 newtons per kilogram). There is negligible or no fuel loss from the solid-core system. The solid-core system, of course, has had the most work done on it. Large-scale tests have been performed on a breadboard engine that has produced specific impulses greater than 700 seconds at thrust levels of about 50,000 pounds (222,000 newtons). The liquid-core reactor would be interesting in the specific impulse range of 1200 to 1500 seconds. Again, any thrust level can be obtained depending on how big or small the reactor is made. The thrust-to-engine weight ratio for these systems would be in the range of 1 to 10. The discouraging feature of the liquid-core system is the high fuel-loss ratio anticipated. Values of 0.01 to 0.1 pound (0.00454 to 0.0454 kilograms) or uranium loss per pound (0.454 kilograms) of hydrogen are expected, if impulses in the range of 1200 to 1500 seconds are desired. The gas-core reactor shows specific impulses in the range of 1500 to 2500 seconds. The thrust levels should be at least as high as the weight so that the thrust-to-weight ratio does not go below 1. Because the engine weight is not expected to be under 100,000 pounds (444,800 newtons), thrust levels higher than 100,000 pounds (448,000 newtons) are of interest. The thrust-to-engine weights, in that case, would run from 1 to 20 pounds per pound (9.8 to 19.7 kilograms). Gas-core reactors tend to be very large, and can have high thrust-to-weight ratios. As in the case of the liquid-core system, the fuel loss that will be attendant with gas cores as envisioned today will be rather high. The loss rates will be 0.01 to 0.1 pound of uranium (0.00454 to 0.0454 kilograms) for each pound (0.454 kilograms) of hydrogen.

Genetic algorithm to optimize the design of main combustor and gas generator in liquid rocket engines

NASA Astrophysics Data System (ADS)

Son, Min; Ko, Sangho; Koo, Jaye

2014-06-01

A genetic algorithm was used to develop optimal design methods for the regenerative cooled combustor and fuel-rich gas generator of a liquid rocket engine. For the combustor design, a chemical equilibrium analysis was applied, and the profile was calculated using Rao's method. One-dimensional heat transfer was assumed along the profile, and cooling channels were designed. For the gas-generator design, non-equilibrium properties were derived from a counterflow analysis, and a vaporization model for the fuel droplet was adopted to calculate residence time. Finally, a genetic algorithm was adopted to optimize the designs. The combustor and gas generator were optimally designed for 30-tonf, 75-tonf, and 150-tonf engines. The optimized combustors demonstrated superior design characteristics when compared with previous non-optimized results. Wall temperatures at the nozzle throat were optimized to satisfy the requirement of 800 K, and specific impulses were maximized. In addition, the target turbine power and a burned-gas temperature of 1000 K were obtained from the optimized gas-generator design.

Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Mitchell, Sonny; Houts, Michael G.; Kim, Tony

2015-01-01

Development efforts in the United States for nuclear thermal propulsion (NTP) systems began with Project Rover (1955-1973) which completed 22 high-power rocket reactor tests. Results indicated that an NTP system with a high thrust-to-weight ratio and a specific impulse greater than 900 s would be feasible. John F. Kennedy, in his historic special address to Congress on the importance of Space on May 25, 1961, said, "First, I believe that this nation should commit itself to achieving the goal, before this decade is out, of landing a man on the Moon and returning him safely to the Earth..." This was accomplished. He also said, "Secondly ... accelerate development of the Rover nuclear rocket. This gives promise of someday providing a means for even more exciting and ambitious exploration of space... to the very end of the solar system itself." The current NTP project focuses on demonstrating the affordability and viability of a fully integrated NTP system with emphasis on fuel fabrication and testing and an affordable development and qualification strategy. The goal is to enable NTP to be considered a mainstream option for supporting human Mars and other missions beyond Earth orbit.

Manufacturing Process Developments for Regeneratively-Cooled Channel Wall Rocket Nozzles

NASA Technical Reports Server (NTRS)

Gradl, Paul; Brandsmeier, Will

2016-01-01

Regeneratively cooled channel wall nozzles incorporate a series of integral coolant channels to contain the coolant to maintain adequate wall temperatures and expand hot gas providing engine thrust and specific impulse. NASA has been evaluating manufacturing techniques targeting large scale channel wall nozzles to support affordability of current and future liquid rocket engine nozzles and thrust chamber assemblies. The development of these large scale manufacturing techniques focus on the liner formation, channel slotting with advanced abrasive water-jet milling techniques and closeout of the coolant channels to replace or augment other cost reduction techniques being evaluated for nozzles. NASA is developing a series of channel closeout techniques including large scale additive manufacturing laser deposition and explosively bonded closeouts. A series of subscale nozzles were completed evaluating these processes. Fabrication of mechanical test and metallography samples, in addition to subscale hardware has focused on Inconel 625, 300 series stainless, aluminum alloys as well as other candidate materials. Evaluations of these techniques are demonstrating potential for significant cost reductions for large scale nozzles and chambers. Hot fire testing is planned using these techniques in the future.

Magnesium and Carbon Dioxide - A Rocket Propellant for Mars Missions

NASA Technical Reports Server (NTRS)

Shafirovich, E. IA.; Shiriaev, A. A.; Goldshleger, U. I.

1993-01-01

A rocket engine for Mars missions is proposed that could utilize CO2 accumulated from the Martian atmosphere as an oxidizer. For use as possible fuel, various metals, their hydrides, and mixtures with hydrogen compounds are considered. Thermodynamic calculations show that beryllium fuels ensure the most impulse but poor inflammability of Be and high toxicity of its compounds put obstacles to their applications. Analysis of the engine performance for other metals together with the parameters of ignition and combustion show that magnesium seems to be the most promising fuel. Ballistic estimates imply that a hopper with the chemical rocket engine on Mg + CO2 propellant could be readily developed. This vehicle would be able to carry out 2-3 ballistic flights on Mars before the final ascent to orbit.

Benchmark of FDNS CFD Code For Direct Connect RBCC Test Data

NASA Technical Reports Server (NTRS)

Ruf, J. H.

2000-01-01

Computational Fluid Dynamics (CFD) analysis results are compared with experimental data from the Pennsylvania State University's (PSU) Propulsion Engineering Research Center (PERC) rocket based combined cycle (RBCC) rocket-ejector experiments. The PERC RBCC experimental hardware was in a direct-connect configuration in diffusion and afterburning (DAB) operation. The objective of the present work was to validate the Finite Difference Navier Stokes (FDNS) CFD code for the rocket-ejector mode internal fluid mechanics and combustion phenomena. A second objective was determine the best application procedures to use FDNS as a predictive/engineering tool. Three-dimensional CFD analysis was performed. Solution methodology and grid requirements are discussed. CFD results are compared to experimental data for static pressure, Raman Spectroscopy species distribution data and RBCC net thrust and specified impulse.

Using Technical Performance Measures

NASA Technical Reports Server (NTRS)

Garrett, Christopher J.; Levack, Daniel J. H.; Rhodes, Russel E.

2011-01-01

All programs have requirements. For these requirements to be met, there must be a means of measurement. A Technical Performance Measure (TPM) is defined to produce a measured quantity that can be compared to the requirement. In practice, the TPM is often expressed as a maximum or minimum and a goal. Example TPMs for a rocket program are: vacuum or sea level specific impulse (lsp), weight, reliability (often expressed as a failure rate), schedule, operability (turn-around time), design and development cost, production cost, and operating cost. Program status is evaluated by comparing the TPMs against specified values of the requirements. During the program many design decisions are made and most of them affect some or all of the TPMs. Often, the same design decision changes some TPMs favorably while affecting other TPMs unfavorably. The problem then becomes how to compare the effects of a design decision on different TPMs. How much failure rate is one second of specific impulse worth? How many days of schedule is one pound of weight worth? In other words, how to compare dissimilar quantities in order to trade and manage the TPMs to meet all requirements. One method that has been used successfully and has a mathematical basis is Utility Analysis. Utility Analysis enables quantitative comparison among dissimilar attributes. It uses a mathematical model that maps decision maker preferences over the tradeable range of each attribute. It is capable of modeling both independent and dependent attributes. Utility Analysis is well supported in the literature on Decision Theory. It has been used at Pratt & Whitney Rocketdyne for internal programs and for contracted work such as the J-2X rocket engine program. This paper describes the construction of TPMs and describes Utility Analysis. It then discusses the use of TPMs in design trades and to manage margin during a program using Utility Analysis.

Comparison of super-high-energy-propulsion-systems based on metallic hydrogen propellant for ES to LEO space transportation

NASA Technical Reports Server (NTRS)

Thierschmann, M.

1990-01-01

The application is studied of metallic H2 as a rocket propellant, which contains a specific energy of about 52 kcal/g in theory yielding a maximum specific impulse of 1700 s. With the convincing advantage of having a density 14 times that of conventional liquid H2/liquid O2 propellants, metallic H2 could satisfy the demands of advanced launch vehicle propulsion for the next millennium. Provided that there is an atomic metallic state of H2, and that this state is metastable at ambient pressure, which still is not proven, the results are given of the study of some important areas, which concern the production of metallic H2, the combustion, chamber cooling, and storage. The results show that the use of metallic H2 as rocket propellant could lead to revolutionary changes in space vehicle philosophy toward small size, small weight, and high performance single stage to orbit systems. The use of high metallic H2 mass fractions results in a dramatic reduction of required propellant volume, while gas temperatures in the combustion chamber exceed 5000 K. Furthermore, it follows, that H2 (liquid or slush) is the most favorable candidate as working fluid. Jet generated noise due to high exhaust velocities could be a problem.

Deimos Methane-Oxygen Rocket Engine Test Results

NASA Astrophysics Data System (ADS)

Engelen, S.; Souverein, L. J.; Twigt, D. J.

This paper presents the results of the first DEIMOS Liquid Methane/Oxygen rocket engine test campaign. DEIMOS is an acronym for `Delft Experimental Methane Oxygen propulsion System'. It is a project performed by students under the auspices of DARE (Delft Aerospace Rocket Engineering). The engine provides a theoretical design thrust of 1800 N and specific impulse of 287 s at a chamber pressure of 40 bar with a total mass flow of 637 g/s. It has links to sustainable development, as the propellants used are one of the most promising so-called `green propellants'-combinations, currently under scrutiny by the industry, and the engine is designed to be reusable. This paper reports results from the provisional tests, which had the aim of verifying the engine's ability to fire, and confirming some of the design assumptions to give confidence for further engine designs. Measurements before and after the tests are used to determine first estimates on feed pressures, propellant mass flows and achieved thrust. These results were rather disappointing from a performance point of view, with an average thrust of a mere 3.8% of the design thrust, but nonetheless were very helpful. The reliability of ignition and stability of combustion are discussed as well. An initial assessment as to the reusability, the flexibility and the adaptability of the engine was made. The data provides insight into (methane/oxygen) engine designs, leading to new ideas for a subsequent design. The ultimate goal of this project is to have an operational rocket and to attempt to set an amateur altitude record.

Laser Ignition Microthruster Experiments on KKS-1

NASA Astrophysics Data System (ADS)

Nakano, Masakatsu; Koizumi, Hiroyuki; Watanabe, Masashi; Arakawa, Yoshihiro

A laser ignition microthruster has been developed for microsatellites. Thruster performances such as impulse and ignition probability were measured, using boron potassium nitrate (B/KNO3) solid propellant ignited by a 1 W CW laser diode. The measured impulses were 60 mNs ± 15 mNs with almost 100 % ignition probability. The effect of the mixture ratios of B/KNO3 on thruster performance was also investigated, and it was shown that mixture ratios between B/KNO3/binder = 28/70/2 and 38/60/2 exhibited both high ignition probability and high impulse. Laser ignition thrusters designed and fabricated based on these data became the first non-conventional microthrusters on the Kouku Kousen Satellite No. 1 (KKS-1) microsatellite that was launched by a H2A rocket as one of six piggyback satellites in January 2009.

The Rationale/Benefits of Nuclear Thermal Rocket Propulsion for NASA's Lunar Space Transportation System

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.

1994-01-01

The solid core nuclear thermal rocket (NTR) represents the next major evolutionary step in propulsion technology. With its attractive operating characteristics, which include high specific impulse (approximately 850-1000 s) and engine thrust-to-weight (approximately 4-20), the NTR can form the basis for an efficient lunar space transportation system (LTS) capable of supporting both piloted and cargo missions. Studies conducted at the NASA Lewis Research Center indicate that an NTR-based LTS could transport a fully-fueled, cargo-laden, lunar excursion vehicle to the Moon, and return it to low Earth orbit (LEO) after mission completion, for less initial mass in LEO than an aerobraked chemical system of the type studied by NASA during its '90-Day Study.' The all-propulsive NTR-powered LTS would also be 'fully reusable' and would have a 'return payload' mass fraction of approximately 23 percent--twice that of the 'partially reusable' aerobraked chemical system. Two NTR technology options are examined--one derived from the graphite-moderated reactor concept developed by NASA and the AEC under the Rover/NERVA (Nuclear Engine for Rocket Vehicle Application) programs, and a second concept, the Particle Bed Reactor (PBR). The paper also summarizes NASA's lunar outpost scenario, compares relative performance provided by different LTS concepts, and discusses important operational issues (e.g., reusability, engine 'end-of life' disposal, etc.) associated with using this important propulsion technology.

Fabrication of High Thermal Conductivity NARloy-Z-Diamond Composite Combustion Chamber Liner for Advanced Rocket Engines

NASA Technical Reports Server (NTRS)

Bhat, Biliyar N.; Greene, Sandra E.; Singh, Jogender

2016-01-01

NARloy-Z alloy (Cu-3 percent, Ag-0.5 percent, Zr) is a state of the art alloy currently used for fabricating rocket engine combustion chamber liners. Research conducted at NASA-MSFC and Penn State – Applied Research Laboratory has shown that thermal conductivity of NARloy-Z can be increased significantly by adding diamonds to form a composite (NARloy-Z-D). NARloy-Z-D is also lighter than NARloy-Z. These attributes make this advanced composite material an ideal candidate for fabricating combustion chamber liner for an advanced rocket engine. Increased thermal conductivity will directly translate into increased turbopump power and increased chamber pressure for improved thrust and specific impulse. This paper describes the process development for fabricating a subscale high thermal conductivity NARloy-Z-D combustion chamber liner using Field Assisted Sintering Technology (FAST). The FAST process uses a mixture of NARloy-Z and diamond powders which is sintered under pressure at elevated temperatures. Several challenges were encountered, i.e., segregation of diamonds, machining the super hard NARloy-Z-D composite, net shape fabrication and nondestructive examination. The paper describes how these challenges were addressed. Diamonds coated with copper (CuD) appear to give the best results. A near net shape subscale combustion chamber liner is being fabricated by diffusion bonding cylindrical rings of NARloy-Z-CuD using the FAST process.

The engineering of a nuclear thermal landing and ascent vehicle utilizing indigenous Martian propellant

NASA Technical Reports Server (NTRS)

Zubrin, Robert M.

1991-01-01

The following paper reports on a design study of a novel space transportation concept known as a 'NIMF' (Nuclear rocket using Indigenous Martian Fuel). The NIMF is a ballistic vehicle which obtains its propellant out of the Martian air by compression and liquefaction of atmospheric CO2. This propellant is subsequently used to generate rocket thrust at a specific impulse of 264 s by being heated to high temperature (2800 K) gas in the NIMFs' nuclear thermal rocket engines. The vehicle is designed to provide surface to orbit and surface to surface transportation, as well as housing, for a crew of three astronauts. It is capable of refueling itself for a flight to its maximum orbit in less than 50 days. The ballistic NIMF has a mass of 44.7 tonnes and, with the assumed 2800 K propellant temperature, is capable of attaining highly energetic (250 km by 34,000 km elliptical) orbits. This allows it to rendezvous with interplanetary transfer vehicles which are only very loosely bound into orbit around Mars. If a propellant temperature of 2000 K is assumed, then low Mars orbit can be attained; while if 3100 K is assumed, then the ballistic NIMF is capable of injecting itself onto a minimum energy transfer orbit to Earth in a direct ascent from the Martian surface.

Powdered aluminum and oxygen rocket propellants: Subscale combustion experiments

NASA Technical Reports Server (NTRS)

Meyer, Mike L.

1993-01-01

Aluminum combined with oxygen has been proposed as a potential lunar in situ propellant for ascent/descent and return missions for future lunar exploration. Engine concepts proposed to use this propellant have not previously been demonstrated, and the impact on performance from combustion and two-phase flow losses could only be estimated. Therefore, combustion tests were performed for aluminum and aluminum/magnesium alloy powders with oxygen in subscale heat-sink rocket engine hardware. The metal powder was pneumatically injected, with a small amount of nitrogen, through the center orifice of a single element O-F-O triplet injector. Gaseous oxygen impinged on the fuel stream. Hot-fire tests of aluminum/oxygen were performed over a mixture ratio range of 0.5 to 3.0, and at a chamber pressure of approximately 480 kPa (70 psia). The theoretical performance of the propellants was analyzed over a mixture ratio range of 0.5 to 5.0. In the theoretical predictions the ideal one-dimensional equilibrium rocket performance was reduced by loss mechanisms including finite rate kinetics, two-dimensional divergence losses, and boundary layer losses. Lower than predicted characteristic velocity and specific impulse performance efficiencies were achieved in the hot-fire tests, and this was attributed to poor mixing of the propellants and two-phase flow effects. Several tests with aluminum/9.8 percent magnesium alloy powder did not indicate any advantage over the pure aluminum fuel.

Laboratory simulation of the rocket motor thrust as a follower force

NASA Technical Reports Server (NTRS)

1990-01-01

Ground tests of solid propellant rocket motors have shown that metal-containing propellants produce various amounts of slag (primarily aluminum oxide), which is trapped in the motor case causing a loss of specific impulse. Although not yet definitely established, the presence of a liquid pool of slag also may contribute to nutational instabilities that have been observed with certain spin-stabilized, upper-stage vehicles. Because of the rocket's axial acceleration - absent in the ground tests - estimates of in-flight slag mass have been very uncertain. Yet such estimates are needed to determine the magnitude of the control authority of the systems required for eliminating the instability. A test rig with an eccentrically mounted hemispherical bowl was designed and built that incorporates a follower force that properly aligns the thrust vector along the axis of spin. A program that computes the motion of a point mass in the spinning and precessing bowl was written. Using various rpm, friction factors, and initial starting conditions, plots were generated showing the trace of the point mass around the inside of the fuel tank. The apparatus will be used extensively during the 1990 to 1991 academic year and incorporate future design features such as a variable nutation angle and a film height measuring instrument. Data obtained on the nutational instability characteristics will be used to determine order-of-magnitude estimates of control authority needed to minimize the sloshing effect.

Laboratory Simulation of the Effect of Rocket Thrust on a Precessing Space Vehicle

NASA Technical Reports Server (NTRS)

Alvarez, Oscar; Bausley, Henry; Cohen, Sam; Falcon-Martin, Miguel; Furumoto, Gary (Editor); Horio, Asikin; Levitt, David; Walsh, Amy

1990-01-01

Ground tests of solid propellant rocket motors have shown that metal-containing propellants produce various amounts of slag (primarily aluminum oxide) which is trapped in the motor case, causing a loss of specific impulse. Although not yet definitely established, the presence of a liquid pool of slag also may contribute to nutational instabilities that have been observed with certain spin-stabilized, upper-stage vehicles. Because of the rocket's axial acceleration, absent in the ground tests, estimates of in-flight slag mass have been very uncertain. Yet such estimates are needed to determine the magnitude of the control authority of the systems required for eliminating the instability. A test rig with an eccentrically mounted hemispherical bowl was designed and built which incorporates a follower force that properly aligns the thrust vector along the axis of spin. A program that computes the motion of a point mass in the spinning and precessing bowl was written. Using various RPMs, friction factors, and initial starting conditions, plots were generated showing the trace of the point mass around the inside of the fuel tank. The apparatus will incorporate future design features such as a variable nutation angle and a film height measuring instrument. Data obtained on the nutational instability characteristics will be used to determine order of magnitude estimates of control authority needed to minimize the sloshing effect.

The Strutjet Rocket Based Combined Cycle Engine

NASA Technical Reports Server (NTRS)

Siebenhaar, A.; Bulman, M. J.; Bonnar, D. K.

1998-01-01

The multi stage chemical rocket has been established over many years as the propulsion System for space transportation vehicles, while, at the same time, there is increasing concern about its continued affordability and rather involved reusability. Two broad approaches to addressing this overall launch cost problem consist in one, the further development of the rocket motor, and two, the use of airbreathing propulsion to the maximum extent possible as a complement to the limited use of a conventional rocket. In both cases, a single-stage-to-orbit (SSTO) vehicle is considered a desirable goal. However, neither the "all-rocket" nor the "all-airbreathing" approach seems realizable and workable in practice without appreciable advances in materials and manufacturing. An affordable system must be reusable with minimal refurbishing on-ground, and large mean time between overhauls, and thus with high margins in design. It has been suggested that one may use different engine cycles, some rocket and others airbreathing, in a combination over a flight trajectory, but this approach does not lead to a converged solution with thrust-to-mass, specific impulse, and other performance and operational characteristics that can be obtained in the different engines. The reason is this type of engine is simply a combination of different engines with no commonality of gas flowpath or components, and therefore tends to have the deficiencies of each of the combined engines. A further development in this approach is a truly combined cycle that incorporates a series of cycles for different modes of propulsion along a flight path with multiple use of a set of components and an essentially single gas flowpath through the engine. This integrated approach is based on realizing the benefits of both a rocket engine and airbreathing engine in various combinations by a systematic functional integration of components in an engine class usually referred to as a rocket-based combined cycle (RBCC) engine. RBCC engines exhibit a high potential for lowering the operating cost of launching payloads into orbit. Two sources of cost reductions can be identified. First, RBCC powered vehicles require only 20% takeoff thrust compared to conventional rockets, thereby lowering the thrust requirements and the replacement cost of the engines. Second, due to the higher structural and thermal margins achievable with RBCC engines coupled with a higher degree of subsystem redundance lower maintenance and operating cost are obtainable.

Solar Electric Propulsion for Mars Exploration

NASA Technical Reports Server (NTRS)

Hack, Kurt J.

1998-01-01

Highly propellant-efficient electric propulsion is being combined with advanced solar power technology to provide a non-nuclear transportation option for the human exploration of Mars. By virtue of its high specific impulse, electric propulsion offers a greater change in spacecraft velocity for each pound of propellant than do conventional chemical rockets. As a result, a mission to Mars based on solar electric propulsion (SEP) would require fewer heavy-lift launches than a traditional all-chemical space propulsion scenario would. Performance, as measured by mass to orbit and trip time, would be comparable to the NASA design reference mission for human Mars exploration, which utilizes nuclear thermal propulsion; but it would avoid the issues surrounding the use of nuclear reactors in space.

An evaluation of composite propulsion for single-stage-to-orbit vehicles designed for horizontal take-off

NASA Technical Reports Server (NTRS)

Martin, J. A.

1977-01-01

Composite propulsion was analyzed for single-stage-to-orbit vehicles designed for horizontal take-off. Trajectory, geometric, and mass analyses were performed to establish the orbital payload capability of six engines. The results indicated that none of the engines performed adequately to deliver payloads to orbit as analyzed. The single-stage turbine and oxidizer-rich gas generator resulted in a low engine specific impulse, and the performance increment of the ejector subsystem was less than that of a separate rocket system with a high combustion pressure. There was a benefit from incorporating a fan into the engine, and removal of the fan from the airstream during the ramjet mode increased the orbital payload capability.

[Bis(imidazolyl)-BH₂]+[bis-(triazolyl)-BH₂]- Ionic Liquids with High Density and Energy Capacity.

PubMed

Jiao, Nianming; Li, Hao; Zhang, Yanqiang; Liu, Long; Zhang, Suojiang

2018-05-15

[Bis(imidazolyl)-BH₂]+[bis(triazolyl)-BH₂]- and [bis- (imidazolyl)-BH₂]+[tris(triazolyl)-BH]- were first synthesized, whose cations and anions were all functionalized with B-H groups and azoles. As B-H groups contributing to hypergolic activity and azole groups improving the energy outputs, the resulting ionic liquids exhibited ignition delay time as low as 20 ms and energy output as high as 461.1 kJ mol-1. Besides, densities (1.07-1.22 g∙cm-3) and density-specific impulse (ρIsp, ~ 360 s g cm-3) reach to relatively high level. It has a great promising for those ionic liquids as sustainable rocket fuels. © 2018 WILEY-VCH Verlag GmbH & Co. KGaA, Weinheim.

SolSTUS: Solar Source Thermal Upper Stage

NASA Astrophysics Data System (ADS)

This paper was written by members of the Utah State University (USU) Space Systems Design class, fall quarter 1993. The class is funded by NASA and administered by the University Space Research Association (USRA). The focus of the class is to give students some experience in design of space systems and as a source of original ideas for NASA. This paper is a summary of the work done by members of the Space Systems Design class during the opening phase of the course. The class was divided into groups to work on different areas of the Solar Thermal Rocket (STR) booster in order to produce a design reference mission that would identify the key design issues. The design reference mission focused upon a small satellite mission to Mars. There are several critical components in a Solar Thermal Rocket. STR's produce a very low thrust, but have a high specific impulse, meaning that they take longer to reach the desired orbit, but use a lot less fuel in doing it. The complexity of the rocket is discussed in this paper. Some of the more critical design problems discussed are: (1) the structural and optical complexity of collecting and focusing sunlight onto a specific point, (2) long term storage of fuel (liquid hydrogen), (3) attitude control while thrusting in an elliptical orbit and orienting the mirrors to collect sunlight, and (4) power and communications for the rocket and it's internal systems. The design reference mission discussed here is a very general mission to Mars. A first order trajectory design has been done and a possible basic science payload for Mars has been suggested. This paper summarizes the design reference mission (DRM) formulated by the USU students during fall quarter and identifies major design challenges that will confront the design team during the next two quarters here at USU.

SolSTUS: Solar Source Thermal Upper Stage

NASA Technical Reports Server (NTRS)

1994-01-01

This paper was written by members of the Utah State University (USU) Space Systems Design class, fall quarter 1993. The class is funded by NASA and administered by the University Space Research Association (USRA). The focus of the class is to give students some experience in design of space systems and as a source of original ideas for NASA. This paper is a summary of the work done by members of the Space Systems Design class during the opening phase of the course. The class was divided into groups to work on different areas of the Solar Thermal Rocket (STR) booster in order to produce a design reference mission that would identify the key design issues. The design reference mission focused upon a small satellite mission to Mars. There are several critical components in a Solar Thermal Rocket. STR's produce a very low thrust, but have a high specific impulse, meaning that they take longer to reach the desired orbit, but use a lot less fuel in doing it. The complexity of the rocket is discussed in this paper. Some of the more critical design problems discussed are: (1) the structural and optical complexity of collecting and focusing sunlight onto a specific point, (2) long term storage of fuel (liquid hydrogen), (3) attitude control while thrusting in an elliptical orbit and orienting the mirrors to collect sunlight, and (4) power and communications for the rocket and it's internal systems. The design reference mission discussed here is a very general mission to Mars. A first order trajectory design has been done and a possible basic science payload for Mars has been suggested. This paper summarizes the design reference mission (DRM) formulated by the USU students during fall quarter and identifies major design challenges that will confront the design team during the next two quarters here at USU.

Expendable solid rocket motor upper stages for the Space Shuttle

NASA Technical Reports Server (NTRS)

Davis, H. P.; Jones, C. M.

1974-01-01

A family of expendable solid rocket motor upper stages has been conceptually defined to provide the payloads for the Space Shuttle with performance capability beyond the low earth operational range of the Shuttle Orbiter. In this concept-feasibility assessment, three new solid rocket motors of fixed impulse are defined for use with payloads requiring levels of higher energy. The conceptual design of these motors is constrained to limit thrusting loads into the payloads and to conserve payload bay length. These motors are combined in various vehicle configurations with stage components derived from other programs for the performance of a broad range of upper-stage missions from spin-stabilized, single-stage transfers to three-axis stabilized, multistage insertions. Estimated payload delivery performance and combined payload mission loading configurations are provided for the upper-stage configurations.

Status of Low Thrust Work at JSC

NASA Technical Reports Server (NTRS)

Condon, Gerald L.

2004-01-01

High performance low thrust (solar electric, nuclear electric, variable specific impulse magnetoplasma rocket) propulsion offers a significant benefit to NASA missions beyond low Earth orbit. As NASA (e.g., Prometheus Project) endeavors to develop these propulsion systems and associated power supplies, it becomes necessary to develop a refined trajectory design capability that will allow engineers to develop future robotic and human mission designs that take advantage of this new technology. This ongoing work addresses development of a trajectory design and optimization tool for assessing low thrust (and other types) trajectories. This work targets to advance the state of the art, enable future NASA missions, enable science drivers, and enhance education. This presentation provides a summary of the low thrust-related JSC activities under the ISP program and specifically, provides a look at a new release of a multi-gravity, multispacecraft trajectory optimization tool (Copernicus) along with analysis performed using this tool over the past year.

Momentum and Heat Flux Measurements in the Exhaust of VASIMR using Helium Propellant

NASA Technical Reports Server (NTRS)

Chavers, D. Gregory; Chang-Diaz, Franklin R.; Irvine, Claude; Squire, Jared P.

2003-01-01

Interplanetary travel requires propulsion systems that can provide high specific impulse (Isp), while also having sufficient thrust to rapidly accelerate large payloads. One such propulsion system is the Variable Specific Impulse Magneto-plasma Rocket (VASIMR), which creates, heats, and ejects plasma to provide variable thrust and Isp, designed to optimally meet the mission requirements. The fraction of the total energy invested in creating the plasma, as compared to the plasma's total kinetic energy, is an important factor in determining the overall system efficiency. In VASIMR, this 'frozen flow loss' is appreciable when at high thrust, but negligible at high Isp. The loss applies to other electric thrusters as well. If some of this energy could be recovered through recombination processes, and reinjected as neutral kinetic energy, the efficiency of VASIMR, in its low Isp/high thrust mode may be improved. An experiment is being conducted to investigate the possibility of recovering some of the energy used to create the plasma by studying the flow characteristics of the charged and neutral particles in the exhaust of the thruster. This paper will cover the measurements of momentum flux and heat flux in the exhaust of the VASIMR test facility using helium as the propellant where the heat flux is comprised of both kinetic and plasma recombination energy. The flux measurements also assist in diagnosing and verifying the plasma conditions in the existing experiment.

XCALIBUR: a Vertical Takeoff TSTO RLV Concept with a HEDM Upperstage and a Scram-Rocket Booster

NASA Astrophysics Data System (ADS)

Bradford, J.

2002-01-01

A new 3rd generation, two-stage-to-orbit (TSTO) reusable launch vehicle (RLV) has been designed. The Xcalibur concept represents a novel approach due to its integration method for the upperstage element of the system. The vertical-takeoff booster, which is powered by rocket-based combined-cycle (RBCC) engines, carries the upperstage internally in the aft section of the airframe to a Mach 15 staging condition. The upperstage is released from the booster and carries the 6,820 kg of payload to low earth orbit (LEO) using its high energy density matter (HEDM) propulsion system. The booster element is capable of returning to the original launch site in a ramjet-cruise propulsion mode. Both the booster and the upperstage utilize advanced technologies including: graphite-epoxy tanks, metal-matrix composites, UHTC TPS materials, electro- mechanical actuators (EMAs), and lightweight subsystems (avionics, power distribution, etc.). The booster system is enabled main propulsion system which utilizes four RBCC engines. These engines operate in four distinct modes: air- augmented rocket (AAR), ramjet, scram-rocket, and all-rocket. The booster operates in AAR mode from takeoff to Mach 3, with ramjet mode operation from Mach 3 to Mach 6. The rocket re-ignition for scram-rocket mode occurs at Mach 6, with all-rocket mode from Mach 14 to the staging condition. The extended utilization of the scram-rocket mode greatly improves vehicle performance by providing superior vehicle acceleration when compared to the scramjet mode performance over the same flight region. Results indicate that the specific impulse penalty due to the scram-rocket mode operation is outweighed by the reduced flight time, smaller vehicle size due to increased mixture ratio, and lower allowable maximum dynamic pressure. A complete vehicle system life-cycle analysis was performed in an automated, multi-disciplinary design environment. Automated disciplinary performance analysis tools include: trajectory (POST), propulsion (SCCREAM), aeroheating (TCAT II), and an Excel spreadsheet for component weight estimation. These tools were automated using `file wrappers' in Phoenix Integration's ModelCenter collaborative design environment. Performance tools utilized for the analysis, but not requiring automation included IDEAS for solid modeling and APAS for the aerodynamic analysis. The paper describes the vehicle concept and operation, discussing the types of technologies used and the nominal flight scenario. A brief discussion explaining the decision-making process for the vehicle configuration is included. For cost predictions, NAFCOM-derived cost estimating relationships were used. Economic predictions were developed using a number of codes, including CABAM (financials), AATe (operations), and GTSafetyII (safety and reliability).

Fast round-trip Mars trajectories

NASA Technical Reports Server (NTRS)

Wilson, Sam

1990-01-01

This paper is concerned with the effect of limiting the overall duration or else the one-way flight time of a round trip to Mars, as reflected in the sum of impulsive velocity increments required of the spacecraft propulsion system. Ignition-to-burnout mass ratios for a hypothetical single stage spacecraft, obtained from the rocket equation by combining these delta-V sums with appropriate values of specific impulse, are used to evaluate the relative effectiveness of four high-thrust propulsion alternatives. If the flight crew goes to the surface of Mars and stays there for the duration of their stopover, it is much cheaper (in terms of delta-V) to minimize their zero-g exposure by limiting the interplanetary transit time of a conjunction-class mission (round trip time = 800-1000 days, Mars stopover = 450-700 days) than to impose the same limit on an opposition-class mission (round trip time less than 600 days, stopover = 40 days). Using solid-core nuclear thermal propulsion to fly a conjunction-class mission, for a moderate mass penalty the interplanetary transit time (each way) probably could be limited to something in the range of 4 to 6 months, depending on the launch year.

Comparison of theoretical and experimental thrust performance of a 1030:1 area ratio rocket nozzle at a chamber pressure of 2413 kN/m2 (350 psia)

NASA Technical Reports Server (NTRS)

Smith, Tamara A.; Pavli, Albert J.; Kacynski, Kenneth J.

1987-01-01

The joint Army. Navy, NASA. Air Force (JANNAF) rocket engine peformnace prediction procedure is based on the use of various reference computer programs. One of the reference programs for nozzle analysis is the Two-Dimensional Kinetics (TDK) Program. The purpose of this report is to calibrate the JANNAF procedure incorporated into the December l984 version of the TDK program for the high-area-ratio rocket engine regime. The calibration was accomplished by modeling the performance of a 1030:1 rocket nozzle tested at NASA Lewis Research Center. A detailed description of the experimental test conditions and TDK input parameters is given. The results show that the computer code predicts delivered vacuum specific impulse to within 0.12 to 1.9 percent of the experimental data. Vacuum thrust coefficient predictions were within + or - 1.3 percent of experimental results. Predictions of wall static pressure were within approximately + or - 5 percent of the measured values. An experimental value for inviscid thrust was obtained for the nozzle extension between area ratios of 427.5 and 1030 by using an integration of the measured wall static pressures. Subtracting the measured thrust gain produced by the nozzle between area ratios of 427.5 and 1030 from the inviscid thrust gain yielded experimental drag decrements of 10.85 and 27.00 N (2.44 and 6.07 lb) for mixture ratios of 3.04 and 4.29, respectively. These values correspond to 0.45 and 1.11 percent of the total vacuum thrust. At a mixture ratio of 4.29, the TDK predicted drag decrement was 16.59 N (3.73 lb), or 0.71 percent of the predicted total vacuum thrust.

Flight Test of the Aerojet 7KS-6000 T-27 Jato Rocket Motor

NASA Technical Reports Server (NTRS)

Bond, Aleck C.; Thibodaux, Joseph G., Jr.

1949-01-01

A flight test of the Aero jet Engineering Corporation's 7KS-6000 T-27 Jato rocket motor was conducted at the Langley Pilotless Aircraft Research Station at Wallops Island, Va, to determine the flight performance characteristics of the motor. The flight test imposed an absolute longitudinal acceleration of 9.8 g upon the rocket motor at 2.8 seconds after launching. The total impulse developed by the motor was 43,400 pound-seconds, and the thrusting time was 7.58 seconds. The maximum thrust was 7200 pounds and occurred at 4.8 seconds after launching. No thrust irregularities attributable to effects of the flight longitudinal acceleration were observed. Certain small thrust irregularities occurred in the flight test which appear to correspond to irregularities observed in static tests conducted elsewhere. A hypothesis regarding the origin of these small irregularities is presented.

Thrust generation experiments on microwave rocket with a beam concentrator for long distance wireless power feeding

NASA Astrophysics Data System (ADS)

Fukunari, Masafumi; Yamaguchi, Toshikazu; Nakamura, Yusuke; Komurasaki, Kimiya; Oda, Yasuhisa; Kajiwara, Ken; Takahashi, Koji; Sakamoto, Keishi

2018-04-01

Experiments using a 1 MW-class gyrotron were conducted to examine a beamed energy propulsion rocket, a microwave rocket with a beam concentrator for long-distance wireless power feeding. The incident beam is transmitted from a beam transmission mirror system. The beam transmission mirror system expands the incident beam diameter to 240 mm to extend the Rayleigh length. The beam concentrator receives the beam and guides it into a 56-mm-diameter cylindrical thruster tube. Plasma ignition and ionization front propagation in the thruster were observed through an acrylic window using a fast-framing camera. Atmospheric air was used as a propellant. Thrust generation was achieved with the beam concentrator. The maximum thrust impulse was estimated as 71 mN s/pulse from a pressure history at the thrust wall at the input energy of 638 J/pulse. The corresponding momentum coupling coefficient, Cm was inferred as 204 N/MW.

Technology Challenges for Deep-Throttle Cryogenic Engines for Space Exploration

NASA Technical Reports Server (NTRS)

Brown, Kendall K.; Nelson, Karl W.

2005-01-01

Historically, cryogenic rocket engines have not been used for in-space applications due to their additional complexity, the mission need for high reliability, and the challenges of propellant boil-off. While the mission and vehicle architectures are not yet defined for the lunar and Martian robotic and human exploration objectives, cryogenic rocket engines offer the potential for higher performance and greater architecture/mission flexibility. In-situ cryogenic propellant production could enable a more robust exploration program by significantly reducing the propellant mass delivered to low earth orbit, thus warranting the evaluation of cryogenic rocket engines versus the hypergolic bi-propellant engines used in the Apollo program. A multi-use engine. one which can provide the functionality that separate engines provided in the Apollo mission architecture, is desirable for lunar and Mars exploration missions because it increases overall architecture effectiveness through commonality and modularity. The engine requirement derivation process must address each unique mission application and each unique phase within each mission. The resulting requirements, such as thrust level, performance, packaging, bum duration, number of operations; required impulses for each trajectory phase; operation after extended space or surface exposure; availability for inspection and maintenance; throttle range for planetary descent, ascent, acceleration limits and many more must be addressed. Within engine system studies, the system and component technology, capability, and risks must be evaluated and a balance between the appropriate amount of technology-push and technology-pull must be addressed. This paper will summarize many of the key technology challenges associated with using high-performance cryogenic liquid propellant rocket engine systems and components in the exploration program architectures. The paper is divided into two areas. The first area describes how the mission requirements affect the engine system requirements and create system level technology challenges. An engine system architecture for multiple applications or a family of engines based upon a set of core technologies, design, and fabrication approaches may reduce overall programmatic cost and risk. The engine system discussion will also address the characterization of engine cycle figures of merit, configurations, and design approaches for some in-space vehicle alternatives under consideration. The second area evaluates the component-level technology challenges induced from the system requirements. Component technology issues are discussed addressing injector, thrust chamber, ignition system, turbopump assembly, and valve design for the challenging requirements of high reliability, robustness, fault tolerance, deep throttling, reasonable performance (with respect to weight and specific impulse).

Technology Challenges for Deep-Throttle Cryogenic Engines for Space Exploration

NASA Astrophysics Data System (ADS)

Brown, Kendall K.; Nelson, Karl W.

2005-02-01

Historically, cryogenic rocket engines have not been used for in-space applications due to their additional complexity, the mission need for high reliability, and the challenges of propellant boil-off. While the mission and vehicle architectures are not yet defined for the lunar and Martian robotic and human exploration objectives, cryogenic rocket engines offer the potential for higher performance and greater architecture/mission flexibility. In-situ cryogenic propellant production could enable a more robust exploration program by significantly reducing the propellant mass delivered to low earth orbit, thus warranting the evaluation of cryogenic rocket engines versus the hypergolic bipropellant engines used in the Apollo program. A multi-use engine, one which can provide the functionality that separate engines provided in the Apollo mission architecture, is desirable for lunar and Mars exploration missions because it increases overall architecture effectiveness through commonality and modularity. The engine requirement derivation process must address each unique mission application and each unique phase within each mission. The resulting requirements, such as thrust level, performance, packaging, burn duration, number of operations; required impulses for each trajectory phase; operation after extended space or surface exposure; availability for inspection and maintenance; throttle range for planetary descent, ascent, acceleration limits and many more must be addressed. Within engine system studies, the system and component technology, capability, and risks must be evaluated and a balance between the appropriate amount of technology-push and technology-pull must be addressed. This paper will summarize many of the key technology challenges associated with using high-performance cryogenic liquid propellant rocket engine systems and components in the exploration program architectures. The paper is divided into two areas. The first area describes how the mission requirements affect the engine system requirements and create system level technology challenges. An engine system architecture for multiple applications or a family of engines based upon a set of core technologies, design, and fabrication approaches may reduce overall programmatic cost and risk. The engine system discussion will also address the characterization of engine cycle figures of merit, configurations, and design approaches for some in-space vehicle alternatives under consideration. The second area evaluates the component-level technology challenges induced from the system requirements. Component technology issues are discussed addressing injector, thrust chamber, ignition system, turbopump assembly, and valve design for the challenging requirements of high reliability, robustness, fault tolerance, deep throttling, reasonable performance (with respect to weight and specific impulse).

Calculation of Supersonic Combustion Using Implicit Schemes

NASA Technical Reports Server (NTRS)

Yoon, Seokkwan; Kwak, Dochan (Technical Monitor)

2003-01-01

One of the technology goals of NASA for advanced space transportation is to develop highly efficient propulsion systems to reduce the cost of payload for space missions. Developments of rockets for the second generation Reusable Launch Vehicle (RLV) in the past several years have been focused on low-cost versions of conventional engines. However, recent changes in the Integrated Space Transportation Program to build a crew transportation vehicle to extend the life of the Space Shuttle fleet might suggest that air-breathing rockets could reemerge as a possible propulsion system for the third generation RLV to replace the Space Shuttle after 2015. The weight of the oxygen tank exceeds thirty percent of the total weight of the Space Shuttle at launch while the payload is only one percent of the total weight. The air-breathing rocket propulsion systems, which consume oxygen in the air, offer clear advantages by making vehicles lighter and more efficient. Experience in the National Aerospace Plane Program in the late 1980s indicates that scramjet engines can achieve high specific impulse for low hypersonic vehicle speeds. Whether taking a form of Rocket Based Combined Cycle (RBCC) or Turbine Based Combined Cycle (TBCC), the scramjet is an essential mode of operation for air-breathing rockets. It is well known that fuel-air mixing and rapid combustion are of crucial importance for the success of scramjet engines since the spreading rate of the supersonic mixing layer decreases as the Mach number increases. A factored form of the Gauss-Seidel relaxation method has been widely used in hypersonic flow research since its first application to non-equilibrium flows. However, difficulties in stability and convergence have been encountered when there is strong interaction between fluid motion and chemical reaction, such as multiple fuel injection problems. The present paper reports the results from investigation of the effect of modifications to the original algorithm on the performance for multiple injectors.

Recent Progress on the VASIMR

NASA Technical Reports Server (NTRS)

ChangDiaz, Franklin R.; Squire, J. P.; Ilin, A. V.; Jacobson, V. T.; Glover, T. W.; Baity, F. W.; Carter, M. D.; Goulding, R. H.; Breizman, B. N.

1999-01-01

Experimental and theoretical studies on the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) have continued through a NASA led collaborative program involving several research groups. In the experimental area, performance characterization of the VASIMR helicon plasma source has been obtained over a portion of the parameter space, with helium and hydrogen propellant. Density (10(exp 18) - 10(exp 19)/ cubic meter) and temperature (5 eV) were measured at moderate degree of ionization in two separate experimental devices. Helicon design improvement and optimization will be discussed. Experiments with the ion cyclotron resonance heating (ICRH) subsection have begun and preliminary results will be discussed. Theoretical picture and integrated numerical simulation continue to be refined to account for the main physics elements of the VASIMR, including RF absorption and particle acceleration with subsequent detachment in the magnetic nozzle.

Mass driver reaction engine characteristics and performance in earth orbital transfer missions

NASA Technical Reports Server (NTRS)

Snow, W. R.; Dunbar, R. S.

1982-01-01

Configurations of a typical mass driver reaction engine (MDRE) are presented and its use for delivery of payloads to geosynchronous orbit (GEO) from low earth orbit (LEO) is discussed. Basic rocket equations are developed for LEO to GEO round-trip missions using a single exhaust velocity. It is shown that exhaust velocities in the 5-10 km/sec range (specific impulse of 500-1000 sec) are well suited for mass drivers, minimizing the overall cost of missions. Payload delivery rate fractions show that there is little to be gained by stretching out LEO to GEO transfer times from 90 to 180 days. It therefore pays to use the shorter trip time, approximately doubling the amount of delivered payload during any fixed time of use of the MDRE.

Experimental investigation of combustor effects on rocket thrust chamber performance

NASA Technical Reports Server (NTRS)

1972-01-01

A design and experimental program to develop special instrumentation systems, design engine hardware, and conduct tests using LOX/GH2 propellants in which the propellant flow stratification was controlled is described. The mixture ratio was varied from 4.6 to 6 overall. The mixture ratios in the core and outer zone were varied from 3.5 to 6 and 5 to 8, respectively. The range in boundary layer coolant was from 0 to 10 percent of the fuel. The nominal chamber pressure and thrust were 225 psia and 7000 pounds, respectively. Pressure and heat flux profiles as well as gas sampling of the exhaust products were obtained. Specific impulse efficiencies of approximately 94 percent and characteristic velocity efficiencies of approximately 97 percent were obtained during the experiments.

Open cycle gas core nuclear rockets

NASA Technical Reports Server (NTRS)

Ragsdale, Robert

1991-01-01

The open cycle gas core engine is a nuclear propulsion device. Propulsion is provided by hot hydrogen which is heated directly by thermal radiation from the nuclear fuel. Critical mass is sustained in the uranium plasma in the center. It has typically 30 to 50 kg of fuel. It is a thermal reactor in the sense that fissions are caused by absorption of thermal neutrons. The fast neutrons go out to an external moderator/reflector material and, by collision, slow down to thermal energy levels, and then come back in and cause fission. The hydrogen propellant is stored in a tank. The advantage of the concept is very high specific impulse because you can take the plasma to any temperature desired by increasing the fission level by withdrawing or turning control rods or control drums.

Europa Sample Return Mission Utilizing High Specific Impulse Propulsion Refueled with Indigenous Resources

NASA Astrophysics Data System (ADS)

Paniagua, J.; Powell, J. R.; Maise, G.

2002-01-01

We have conducted studies of a revolutionary new concept for conducting a Europa Sample Return Mission. Robotic spacecraft exploration of the Solar System has been severely constrained by the large energy requirements of interplanetary trajectories and the inherent delta V limitations of chemical rockets. Current missions use gravitational assists from intermediate planets to achieve these high-energy trajectories restricting payload size and increasing flight times. We propose a 6-year Europa Sample Return mission with very modest launch requirements enabled by MITEE. A new nuclear thermal propulsion engine design, termed MITEE (MIniature reacTor EnginE), has over twice the delta V capability of H2/O2 rockets (and much greater when refueled with H2 propellant from indigenous extraterrestrial resources) enabling unique missions that are not feasible with chemical propulsion. The MITEE engine is a compact, ultra-lightweight, thermal nuclear rocket that uses hydrogen as the propellant. MITEE, with its small size (50 cm O.D.), low mass (200 kg), and high specific impulse (~1000 sec), can provide a quantum leap in the capability for space science and exploration missions. The Robotic Europa Explorer (REE) spacecraft has a two-year outbound direct trajectory and lands on the satellite surface for an approximate 9 month stay. During this time, the vehicle is refueled with H2 propellant derived from Europa ice by the Autonomous Propellant Producer (APP), while collecting samples and searching for life. A small nuclear-heated submarine probe, the Autonomous Submarine Vehicle (ASV), based on MITEE technology, would melt through the ice and explore the undersea realm. The spacecraft has approximately a three year return to Earth after departure from Europa with samples onboard. Spacecraft payload is 430 kg at the start of the mission and can be launched with a single, conventional medium-sized Delta III booster. The spacecraft can bring back 25 kg of samples from Europa. Europa, in the Jovian system, is a high priority target for an outer Solar System exploration mission. More than a decade ago the Voyager spacecraft revealed Europa as a world swathed in ice and geologically young. NASA's Galileo spacecraft passed approximately 500 miles above the surface and provided detailed images of Europa's terrain marked by a dynamic topology that appeared to be remnants of ice volcanoes or geysers. The surface temperature averages a chilly -200° C. The pictures appear to show a relatively young surface of ice, possibly only 1 km thick in some places. Internal heating of Europa from Jupiter's tidal pull could form an ocean of liquid water beneath the surface. More recently, Ganymede and Callisto are believed to be ocean-bearing Jovian moons based on magnetometer measurements from the Galileo spacecraft. If liquid water exists, life may also. NASA plans to send an orbiting spacecraft to Europa to measure the thickness of the ice and to detect if an underlying liquid ocean exists. This mission would precede the proposed Europa Sample Return mission, which includes dispatching an autonomous submarine-like vehicle that could melt through the ice and explore the undersea realm. Because of the large energy requirements typical of these ambitious solar system science missions, use of chemical rockets results in interplanetary spacecraft that are prohibitive in terms of Initial Mass in Low- Earth Orbit (IMLEO) and cost. For example, using chemical rockets to return samples from Europa appears to be technically impractical, as it would require large delta V and launch vehicle capabilities. On the other hand, use of nuclear thermal rockets will significantly reduce IMLEO and, subsequently, costs. Moreover, nuclear thermal rockets can utilize extraterrestrial resources as propellants, an option not practical with chemical rockets. This "refueling" capability would enable nuclear rockets to carry out very high-energy missions, such as the return of large amounts of extraterrestrial material to Earth. The Europa missions considered in this proposal will be restricted to starting from LEO only after being placed in a stable orbit by a launch vehicle. This simplifies and eases the safety issues and mitigates political concerns. High propulsive efficiency of the MITEE engine yields the benefits of reduced transit time and a smaller launch vehicle.

Analytical study of nozzle performance for nuclear thermal rockets

NASA Technical Reports Server (NTRS)

Davidian, Kenneth O.; Kacynski, Kenneth J.

1991-01-01

Nuclear propulsion has been identified as one of the key technologies needed for human exploration of the Moon and Mars. The Nuclear Thermal Rocket (NTR) uses a nuclear reactor to heat hydrogen to a high temperature followed by expansion through a conventional convergent-divergent nozzle. A parametric study of NTR nozzles was performed using the Rocket Engine Design Expert System (REDES) at the NASA Lewis Research Center. The REDES used the JANNAF standard rigorous methodology to determine nozzle performance over a range of chamber temperatures, chamber pressures, thrust levels, and different nozzle configurations. A design condition was set by fixing the propulsion system exit radius at five meters and throat radius was varied to achieve a target thrust level. An adiabatic wall was assumed for the nozzle, and its length was assumed to be 80 percent of a 15 degree cone. The results conclude that although the performance of the NTR, based on infinite reaction rates, looks promising at low chamber pressures, finite rate chemical reactions will cause the actual performance to be considerably lower. Parameters which have a major influence on the delivered specific impulse value include the chamber temperature and the chamber pressures in the high thrust domain. Other parameters, such as 2-D and boundary layer effects, kinetic rates, and number of nozzles, affect the deliverable performance of an NTR nozzle to a lesser degree. For a single nozzle, maximum performance of 930 seconds and 1030 seconds occur at chamber temperatures of 2700 and 3100 K, respectively.

Gas Core Nuclear Rocket Feasibility Project

NASA Technical Reports Server (NTRS)

Howe, S. D.; DeVolder, B.; Thode, L.; Zerkle, D.

1997-01-01

The next giant leap for mankind will be the human exploration of Mars. Almost certainly within the next thirty years, a human crew will brave the isolation, the radiation, and the lack of gravity to walk on and explore the Red planet. However, because the mission distances and duration will be hundreds of times greater than the lunar missions, a human crew will face much greater obstacles and a higher risk than those experienced during the Apollo program. A single solution to many of these obstacles is to dramatically decrease the mission duration by developing a high performance propulsion system. The gas core nuclear rocket (GCNR) has the potential to be such a system. The gas core concept relies on the use of fluid dynamic forces to create and maintain a vortex. The vortex is composed of a fissile material which will achieve criticality and produce high power levels. By radiatively coupling to the surrounding fluids, extremely high temperatures in the propellant and, thus, high specific impulses can be generated. The ship velocities enabled by such performance may allow a 9 month round trip, manned Mars mission to be considered. Alternatively, one might consider slightly longer missions in ships that are heavily shielded against the intense Galactic Cosmic Ray flux to further reduce the radiation dose to the crew. The current status of the research program at the Los Alamos National Laboratory into the gas core nuclear rocket feasibility will be discussed.

Computer program for prediction of fuel consumption statistical data for an upper stage three-axes stabilized on-off control system

NASA Technical Reports Server (NTRS)

1982-01-01

A FORTRAN coded computer program and method to predict the reaction control fuel consumption statistics for a three axis stabilized rocket vehicle upper stage is described. A Monte Carlo approach is used which is more efficient by using closed form estimates of impulses. The effects of rocket motor thrust misalignment, static unbalance, aerodynamic disturbances, and deviations in trajectory, mass properties and control system characteristics are included. This routine can be applied to many types of on-off reaction controlled vehicles. The pseudorandom number generation and statistical analyses subroutines including the output histograms can be used for other Monte Carlo analyses problems.

Nuclear Thermal Propulsion (NTP): A Proven Growth Technology for Human NEO/Mars Exploration Missions

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; McCurdy, David R.; Packard, Thomas W.

2012-01-01

The nuclear thermal rocket (NTR) represents the next "evolutionary step" in high performance rocket propulsion. Unlike conventional chemical rockets that produce their energy through combustion, the NTR derives its energy from fission of Uranium-235 atoms contained within fuel elements that comprise the engine s reactor core. Using an "expander" cycle for turbopump drive power, hydrogen propellant is raised to a high pressure and pumped through coolant channels in the fuel elements where it is superheated then expanded out a supersonic nozzle to generate high thrust. By using hydrogen for both the reactor coolant and propellant, the NTR can achieve specific impulse (Isp) values of 900 seconds (s) or more - twice that of today s best chemical rockets. From 1955 - 1972, twenty rocket reactors were designed, built and ground tested in the Rover and NERVA (Nuclear Engine for Rocket Vehicle Applications) programs. These programs demonstrated: (1) high temperature carbide-based nuclear fuels; (2) a wide range of thrust levels; (3) sustained engine operation; (4) accumulated lifetime at full power; and (5) restart capability - all the requirements needed for a human Mars mission. Ceramic metal "cermet" fuel was pursued as well, as a backup option. The NTR also has significant "evolution and growth" capability. Configured as a "bimodal" system, it can generate its own electrical power to support spacecraft operational needs. Adding an oxygen "afterburner" nozzle introduces a variable thrust and Isp capability and allows bipropellant operation. In NASA s recent Mars Design Reference Architecture (DRA) 5.0 study, the NTR was selected as the preferred propulsion option because of its proven technology, higher performance, lower launch mass, versatile vehicle design, simple assembly, and growth potential. In contrast to other advanced propulsion options, no large technology scale-ups are required for NTP either. In fact, the smallest engine tested during the Rover program - the 25,000 lbf (25 klbf) "Pewee" engine is sufficient when used in a clustered engine arrangement. The "Copernicus" crewed spacecraft design developed in DRA 5.0 has significant capability and a human exploration strategy is outlined here that uses Copernicus and its key components for precursor near Earth object (NEO) and Mars orbital missions prior to a Mars landing mission. The paper also discusses NASA s current activities and future plans for NTP development that include system-level Technology Demonstrations - specifically ground testing a small, scalable NTR by 2020, with a flight test shortly thereafter.

JPRS Report, Science & Technology, China

DTIC Science & Technology

1991-10-22

ZHONGGUO KEXUE BAO, 30 Aug 91] .......................................... 22 Shanghai Scientist Develops State-of-the-Art Liquid-Crystal Light Valve...the angle of attack will gradu- direction of the final velocity vector of the satellite are ally decrease under the action of aerodynamic moments...impulse and the direction of the thrust vector of the The recovery system, is located inside the sealed reentry retro-rocket engine, errors in the

Handbook of recommended practices for the determination of liquid monopropellant rocket engine performance

NASA Technical Reports Server (NTRS)

Bjorklund, R. A.; Rogero, R. S.; Baerwald, R. K.

1979-01-01

The design, installation, and operation of systems to be used for directly measuring quantities of fundamental importance to the determination of monopropellant thruster performance is described. Areas covered include: (1) force and impulse measurement; (2) propellant mass usage and flow measurement; (3) pressure measurement; (4) temperature measurement; (5) exhaust gas composition measurement; and (6) data reduction and performance determination.

Hard X-ray Detectability of Small Impulsive Heating Events in the Solar Corona

NASA Astrophysics Data System (ADS)

Glesener, L.; Klimchuk, J. A.; Bradshaw, S. J.; Marsh, A.; Krucker, S.; Christe, S.

2015-12-01

Impulsive heating events ("nanoflares") are a candidate to supply the solar corona with its ~2 MK temperature. These transient events can be studied using extreme ultraviolet and soft X-ray observations, among others. However, the impulsive events may occur in tenuous loops on small enough timescales that the heating is essentially not observed due to ionization timescales, and only the cooling phase is observed. Bremsstrahlung hard X-rays could serve as a more direct and prompt indicator of transient heating events. A hard X-ray spacecraft based on the direct-focusing technology pioneered by the Focusing Optics X-ray Solar Imager (FOXSI) sounding rocket could search for these direct signatures. In this work, we use the hydrodynamical EBTEL code to simulate differential emission measures produced by individual heating events and by ensembles of such events. We then directly predict hard X-ray spectra and consider their observability by a future spaceborne FOXSI, and also by the RHESSI and NuSTAR spacecraft.

Performance Optimization of the Gasdynamic Mirror Propulsion System

NASA Technical Reports Server (NTRS)

Emrich, William J., Jr.; Kammash, Terry

1999-01-01

Nuclear fusion appears to be a most promising concept for producing extremely high specific impulse rocket engines. Engines such as these would effectively open up the solar system to human exploration and would virtually eliminate launch window restrictions. A preliminary vehicle sizing and mission study was performed based on the conceptual design of a Gasdynamic Mirror (GDM) fusion propulsion system. This study indicated that the potential specific impulse for this engine is approximately 142,000 sec. with about 22,100 N of thrust using a deuterium-tritium fuel cycle. The engine weight inclusive of the power conversion system was optimized around an allowable engine mass of 1500 Mg assuming advanced superconducting magnets and a Field Reversed Configuration (FRC) end plug at the mirrors. The vehicle habitat, lander, and structural weights are based on a NASA Mars mission study which assumes the use of nuclear thermal propulsion' Several manned missions to various planets were analyzed to determine fuel requirements and launch windows. For all fusion propulsion cases studied, the fuel weight remained a minor component of the total system weight regardless of when the missions commenced. In other words, the use of fusion propulsion virtually eliminates all mission window constraints and effectively allows unlimited manned exploration of the entire solar system. It also mitigates the need to have a large space infrastructure which would be required to support the transfer of massive amounts of fuel and supplies to lower a performing spacecraft.

Next Generation H- Ion Sources for the SNS

NASA Astrophysics Data System (ADS)

Welton, R. F.; Stockli, M. P.; Murray, S. N.; Crisp, D.; Carmichael, J.; Goulding, R. H.; Han, B.; Tarvainen, O.; Pennisi, T.; Santana, M.

2009-03-01

The U.S. Spallation Neutron Source (SNS) is the leading accelerator-based, pulsed neutron-scattering facility, currently in the process of ramping up neutron production. In order to insure meeting operational requirements as well as providing for future facility beam power upgrades, a multifaceted H- ion source development program is ongoing. This work discusses several aspects of this program, specifically the design and first beam measurements of an RF-driven, external antenna H- ion source based on an A1N ceramic plasma chamber, elemental and chromate Cs-systems, and plasma ignition gun. Unanalyzed beam currents of up to ˜100 mA (60 Hz, 1 ms) have been observed and sustained currents >60 mA (60 Hz, 1 ms) have been demonstrated on the test stand. Accelerated beam currents of ˜40 mA have also been demonstrated into the SNS front end. Data are also presented describing the first H- beam extraction experiments from a helicon plasma generator based on the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) engine design.

Helicon Plasma Source Optimization Studies for VASIMR

NASA Technical Reports Server (NTRS)

Goulding, R. H.; Baity, F. W.; Barber, G. C.; Carter, M. D.; ChangDiaz, F. R.; Pavarin, D.; Sparks, D. O.; Squire J. P.

1999-01-01

A helicon plasma source at Oak Ridge National Laboratory is being used to investigate operating scenarios relevant to the VASIMR (VAriable Specific Impulse Magnetoplasma Rocket). These include operation at high magnetic field (> = 0.4 T), high frequency (<= 30 MHz), high power (< = 3 kW), and with light ions (He+, H+). To date, He plasmas have been produced with n(sub e0) = 1.7 x 10(exp 19)/cu m (measured with an axially movable 4mm microwave interferometer), with Pin = I kW at f = 13.56 MHz and absolute value of B(sub 0) = 0.16 T. In the near future, diagnostics including a mass flow meter and a gridded energy analyzer array will be added to investigate fueling efficiency and the source power balance. The latest results, together with modeling results using the EMIR rf code, will be presented.

Advanced APS impacts on vehicle payloads

NASA Technical Reports Server (NTRS)

Schneider, Steven J.; Reed, Brian D.

1989-01-01

Advanced auxiliary propulsion system (APS) technology has the potential to both, increase the payload capability of earth-to-orbit (ETO) vehicles by reducing APS propellant mass, and simplify ground operations and logistics by reducing the number of fluids on the vehicle and eliminating toxic, corrosive propellants. The impact of integrated cryogenic APS on vehicle payloads is addressed. In this system, launch propulsion system residuals are scavenged from integral launch propulsion tanks for use in the APS. Sufficient propellant is preloaded into the APS to return to earth with margin and noncomplete scavenging assumed. No propellant conditioning is required by the APS, but ambient heat soak is accommodated. High temperature rocket materials enable the use of the unconditioned hydrogen/oxygen in the APS and are estimated to give APS rockets specific impulse of up to about 444 sec. The payload benefits are quantified and compared with an uprated monomethylhydrazine/nitrogen tetroxide system in a conservative fashion, by assuming a 25.5 percent weight growth for the hydrogen/oxygen system and a 0 percent weight growth for the uprated system. The combination of scavenging and high performance gives payload impacts which are highly mission specific. A payload benefit of 861 kg (1898 lbm) was estimated for a Space Station Freedom rendezvous mission and 2099 kg (4626 lbm) for a sortie mission, with payload impacts varying with the amount of launch propulsion residual propellants. Missions without liquid propellant scavenging were estimated to have payload penalties, however, operational benefits were still possible.

Advanced APS Impacts on Vehicle Payloads

NASA Technical Reports Server (NTRS)

Schneider, Steven J.; Reed, Brian D.

1989-01-01

Advanced auxiliary propulsion system (APS) technology has the potential to both, increase the payload capability of earth-to-orbit (ETO) vehicles by reducing APS propellant mass, and simplify ground operations and logistics by reducing the number of fluids on the vehicle and eliminating toxic, corrosive propellants. The impact of integrated cryogenic APS on vehicle payloads is addressed. In this system, launch propulsion system residuals are scavenged from integral launch propulsion tanks for use in the APS. Sufficient propellant is preloaded into the APS to return to earth with margin and noncomplete scavenging assumed. No propellant conditioning is required by the APS, but ambient heat soak is accommodated. High temperature rocket materials enable the use of the unconditioned hydrogen/oxygen in the APS and are estimated to give APS rockets specific impulse of up to about 444 sec. The payload benefits are quantified and compared with an uprated monomethyl hydrazine/nitrogen tetroxide system in a conservative fashion, by assuming a 25.5 percent weight growth for the hydrogen/oxygen system and a 0 percent weight growth for the uprated system. The combination and scavenging and high performance gives payload impacts which are highly mission specific. A payload benefit of 861 kg (1898 lbm) was estimated for a Space Station Freedom rendezvous mission and 2099 kg (4626 lbm) for a sortie mission, with payload impacts varying with the amount of launch propulsion residual propellants. Missions without liquid propellant scavenging were estimated to have payload penalties, however, operational benefits were still possible.

Homotopy method for optimization of variable-specific-impulse low-thrust trajectories

NASA Astrophysics Data System (ADS)

Chi, Zhemin; Yang, Hongwei; Chen, Shiyu; Li, Junfeng

2017-11-01

The homotopy method has been used as a useful tool in solving fuel-optimal trajectories with constant-specific-impulse low thrust. However, the specific impulse is often variable for many practical solar electric power-limited thrusters. This paper investigates the application of the homotopy method for optimization of variable-specific-impulse low-thrust trajectories. Difficulties arise when the two commonly-used homotopy functions are employed for trajectory optimization. The optimal power throttle level and the optimal specific impulse are coupled with the commonly-used quadratic and logarithmic homotopy functions. To overcome these difficulties, a modified logarithmic homotopy function is proposed to serve as a gateway for trajectory optimization, leading to decoupled expressions of both the optimal power throttle level and the optimal specific impulse. The homotopy method based on this homotopy function is proposed. Numerical simulations validate the feasibility and high efficiency of the proposed method.

Unique aspects of impulsive traits in substance use and overeating: specific contributions of common assessments of impulsivity.

PubMed

Beaton, Derek; Abdi, Hervé; Filbey, Francesca M

2014-11-01

Abstract Background: Impulsivity is a complex trait often studied in substance abuse and overeating disorders, but the exact nature of impulsivity traits and their contribution to these disorders are still debated. Thus, understanding how to measure impulsivity is essential for comprehending addictive behaviors. Identify unique impulsivity traits specific to substance use and overeating. Impulsive Sensation Seeking (ImpSS) and Barratt's Impulsivity scales (BIS) Scales were analyzed with a non-parametric factor analytic technique (discriminant correspondence analysis) to identify group-specific traits on 297 individuals from five groups: Marijuana (n = 88), Nicotine (n = 82), Overeaters (n = 27), Marijuauna + Nicotine (n = 63), and CONTROLs (n = 37). A significant overall factor structure revealed three components of impulsivity that explained respectively 50.19% (pperm < 0.0005), 24.18% (pperm < 0.0005), and 15.98% (pperm < 0.0005) of the variance. All groups were significantly different from one another. When analyzed together, the BIS and ImpSS produce a multi-factorial structure that identified the impulsivity traits specific to these groups. The group specific traits are (1) CONTROL: low impulse, avoids thrill-seeking behaviors; (2) Marijuana: seeks mild sensation, is focused and attentive; (3) Marijuana + Nicotine: pursues thrill-seeking, lacks focus and attention; (4) Nicotine: lacks focus and planning; (5) Overeating: lacks focus, but plans (short and long term). Our results reveal impulsivity traits specific to each group. This may provide better criteria to define spectrums and trajectories - instead of categories - of symptoms for substance use and eating disorders. Defining symptomatic spectrums could be an important step forward in diagnostic strategies.

Unique aspects of impulsive traits in substance use and overeating: specific contributions of common assessments of impulsivity

PubMed Central

Beaton, Derek; Abdi, Hervé; Filbey, Francesca M.

2015-01-01

Background Impulsivity is a complex trait often studied in substance abuse and overeating disorders, but the exact nature of impulsivity traits and their contribution to these disorders are still debated. Thus, understanding how to measure impulsivity is essential for comprehending addictive behaviors. Objectives Identify unique impulsivity traits specific to substance use and overeating. Methods Impulsive Sensation Seeking (ImpSS) and Barratt’s Impulsivity scales (BIS) Scales were analyzed with a non-parametric factor analytic technique (discriminant correspondence analysis) to identify group-specific traits on 297 individuals from five groups: Marijuana (n = 88), Nicotine (n = 82), Overeaters (n = 27), Marijuauna + Nicotine (n = 63), and Controls (n = 37). Results A significant overall factor structure revealed three components of impulsivity that explained respectively 50.19% (pperm<0.0005), 24.18% (pperm<0.0005), and 15.98% (pperm<0.0005) of the variance. All groups were significantly different from one another. When analyzed together, the BIS and ImpSS produce a multi-factorial structure that identified the impulsivity traits specific to these groups. The group specific traits are (1) Control: low impulse, avoids thrill-seeking behaviors; (2) Marijuana: seeks mild sensation, is focused and attentive; (3) Marijuana + Nicotine: pursues thrill-seeking, lacks focus and attention; (4) Nicotine: lacks focus and planning; (5) Overeating: lacks focus, but plans (short and long term). Conclusions Our results reveal impulsivity traits specific to each group. This may provide better criteria to define spectrums and trajectories – instead of categories – of symptoms for substance use and eating disorders. Defining symptomatic spectrums could be an important step forward in diagnostic strategies. PMID:25115831

Feasibility Study of SSTO Base Heating Simulation in Pulsed-Type Facilities

NASA Technical Reports Server (NTRS)

Park, Chung Sik; Sharma, Surendra; Edwards, Thomas A. (Technical Monitor)

1995-01-01

A laboratory simulation of the base heating environment of the proposed reusable Single-Stage-To-Orbit vehicle during its ascent flight was proposed. The rocket engine produces CO2 and H2, which are the main combustible components of the exhaust effluent. The burning of these species, known as afterburning, enhances the base region gas temperature as well as the base heating. To determine the heat flux on the SSTO vehicle, current simulation focuses on the thermochemistry of the afterburning, thermophysical properties of the base region gas, and ensuing radiation from the gas. By extrapolating from the Saturn flight data, the Damkohler number for the afterburning of SSTO vehicle is estimated to be of the order of 10. The limitations on the material strengths limit the laboratory simulation of the flight Damkohler number as well as other flow parameters. A plan is presented in impulse facilities using miniature rocket engines which generate the simulated rocket plume by electric ally-heating a H2/CO2 mixture.

NTREES Testing and Operations Status

NASA Technical Reports Server (NTRS)

Emrich, Bill

2007-01-01

Nuclear Thermal Rockets or NTR's have been suggested as a propulsion system option for vehicles traveling to the moon or Mars. These engines are capable of providing high thrust at specific impulses at least twice that of today's best chemical engines. The performance constraints on these engines are mainly the result of temperature limitations on the fuel coupled with a limited ability to withstand chemical attack by the hot hydrogen propellant. To operate at maximum efficiency, fuel forms are desired which can withstand the extremely hot, hostile environment characteristic of NTR operation for at least several hours. The simulation of such an environment would require an experimental device which could simultaneously approximate the power, flow, and temperature conditions which a nuclear fuel element (or partial element) would encounter during NTR operation. Such a simulation would allow detailed studies of the fuel behavior and hydrogen flow characteristics under reactor like conditions to be performed. Currently, the construction of such a simulator has been completed at the Marshall Space Flight Center, and will be used in the future to evaluate a wide variety of fuel element designs and the materials of which they are fabricated. This present work addresses the operational status of the Nuclear Thermal Rocket Element Environmental Simulator or NTREES and some of the design considerations which were considered prior to and during its construction.

Parametric investigation of secondary injection in post-chamber on combustion performance for hybrid rocket motor

NASA Astrophysics Data System (ADS)

Cai, Guobiao; Cao, Binbin; Zhu, Hao; Tian, Hui; Ma, Xuan

2017-11-01

The objective of this effort is to study the combustion performance of a hybrid rocket motor with the help of 3D steady-state numerical simulation, which applies 90% hydrogen peroxide as the oxidizer and hydroxyl-terminated polybutadiene as the fuel. A method of secondary oxidizer injection in post-chamber is introduced to investigate the flow field characteristics and combustion efficiency. The secondary injection medium is the mixed gas coming from liquid hydrogen peroxide catalytic decomposition. The secondary injectors are uniformly set along the circumferential direction of the post-chamber. The simulation results obtained by above model are verified by experimental data. Three influencing parameters are considered: secondary injection diameter, secondary injection angle and secondary injection numbers. Simulation results reveals that this design could improve the combustion efficiency with respect to the same motor without secondary injection. Besides, the secondary injection almost has no effect on the regression rate and fuel sueface temperature distribution. It is also presented that the oxidizer is injected by 8 secondary injectors with a diameter of 7-8 mm in the direction of 120°in post-chamber is identified as the optimized secondary injection pattern, through which combustion efficiency, specific impulse efficiency as well as utilization of propellants are all improved obviously.

Fuels and Space Propellants for Reusable Launch Vehicles: A Small Business Innovation Research Topic and Its Commercial Vision

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan A.

1997-01-01

Under its Small Business Innovation Research (SBIR) program (and with NASA Headquarters support), the NASA Lewis Research Center has initiated a topic entitled "Fuels and Space Propellants for Reusable Launch Vehicles." The aim of this project would be to assist in demonstrating and then commercializing new rocket propellants that are safer and more environmentally sound and that make space operations easier. Soon it will be possible to commercialize many new propellants and their related component technologies because of the large investments being made throughout the Government in rocket propellants and the technologies for using them. This article discusses the commercial vision for these fuels and propellants, the potential for these propellants to reduce space access costs, the options for commercial development, and the benefits to nonaerospace industries. This SBIR topic is designed to foster the development of propellants that provide improved safety, less environmental impact, higher density, higher I(sub sp), and simpler vehicle operations. In the development of aeronautics and space technology, there have been limits to vehicle performance imposed by traditionally used propellants and fuels. Increases in performance are possible with either increased propellant specific impulse, increased density, or both. Flight system safety will also be increased by the use of denser, more viscous propellants and fuels.

Compact Fuel Element Environment Test

NASA Technical Reports Server (NTRS)

Bradley, D. E.; Mireles, O. R.; Hickman, R. R.; Broadway, J. W.

2012-01-01

Deep space missions with large payloads require high specific impulse (I(sub sp)) and relatively high thrust to achieve mission goals in reasonable time frames. Conventional, storable propellants produce average I(sub sp). Nuclear thermal rockets (NTRs) capable of high I(sub sp) thrust have been proposed. NTR employs heat produced by fission reaction to heat and therefore accelerate hydrogen, which is then forced through a rocket nozzle providing thrust. Fuel element temperatures are very high (up to 3,000 K) and hydrogen is highly reactive with most materials at high temperatures. Data covering the effects of high-temperature hydrogen exposure on fuel elements are limited. The primary concern is the mechanical failure of fuel elements that employ high melting point metals, ceramics, or a combination (cermet) as a structural matrix into which the nuclear fuel is distributed. It is not necessary to include fissile material in test samples intended to explore high-temperature hydrogen exposure of the structural support matrices. A small-scale test bed designed to heat fuel element samples via noncontact radio frequency heating and expose samples to hydrogen for typical mission durations has been developed to assist in optimal material and manufacturing process selection without employing fissile material. This Technical Memorandum details the test bed design and results of testing conducted to date.

Thermal-hydraulics Analysis of a Radioisotope-powered Mars Hopper Propulsion System

DOE Office of Scientific and Technical Information (OSTI.GOV)

Robert C. O'Brien; Andrew C. Klein; William T. Taitano

Thermal-hydraulics analyses results produced using a combined suite of computational design and analysis codes are presented for the preliminary design of a concept Radioisotope Thermal Rocket (RTR) propulsion system. Modeling of the transient heating and steady state temperatures of the system is presented. Simulation results for propellant blow down during impulsive operation are also presented. The results from this study validate the feasibility of a practical thermally capacitive RTR propulsion system.

Internal Flow Analysis of Large L/D Solid Rocket Motors

NASA Technical Reports Server (NTRS)

Laubacher, Brian A.

2000-01-01

Traditionally, Solid Rocket Motor (SRM) internal ballistic performance has been analyzed and predicted with either zero-dimensional (volume filling) codes or one-dimensional ballistics codes. One dimensional simulation of SRM performance is only necessary for ignition modeling, or for motors that have large length to port diameter ratios which exhibit an axial "pressure drop" during the early burn times. This type of prediction works quite well for many types of motors, however, when motor aspect ratios get large, and port to throat ratios get closer to one, two dimensional effects can become significant. The initial propellant grain configuration for the Space Shuttle Reusable Solid Rocket Motor (RSRM) was analyzed with 2-D, steady, axi-symmetric computational fluid dynamics (CFD). The results of the CFD analysis show that the steady-state performance prediction at the initial burn geometry, in general, agrees well with 1-D transient prediction results at an early time, however, significant features of the 2-D flow are captured with the CFD results that would otherwise go unnoticed. Capturing these subtle differences gives a greater confidence to modeling accuracy, and additional insight with which to model secondary internal flow effects like erosive burning. Detailed analysis of the 2-D flowfield has led to the discovery of its hidden 1-D isentropic behavior, and provided the means for a thorough and simplified understanding of internal solid rocket motor flow. Performance parameters such as nozzle stagnation pressure, static pressure drop, characteristic velocity, thrust and specific impulse are discussed in detail and compared for different modeling and prediction methods. The predicted performance using both the 1-D codes and the CFD results are compared with measured data obtained from static tests of the RSRM. The differences and limitations of predictions using ID and 2-D flow fields are discussed and some suggestions for the design of large L/D motors and more critically, motors with port to throat ratios near one, are covered.

Design analysis and risk assessment for a single stage to orbit nuclear thermal rocket

NASA Astrophysics Data System (ADS)

Labib, Satira I.

Recent advances in high power density fuel materials have renewed interest in nuclear thermal rockets (NTRs) as a viable propulsion technology for future space exploration. This thesis describes the design of three NTR reactor engines designed for the single stage to orbit launch of payloads from 1-15 metric tons. Thermal hydraulic and rocket engine analyses indicate that the proposed rocket engines are able to reach specific impulses in excess of 700 seconds. Neutronics analyses performed using MCNP5 demonstrate that the hot excess reactivity, shutdown margin, and submersion criticality requirements are satisfied for each NTR reactor. The reactors each consist of a 40 cm diameter core packed with hexagonal tungsten cermet fuel elements. The core is surrounded by radial and axial beryllium reflectors and eight boron carbide control drums. At the same power level, the 40 cm reactor results in the lowest radiation dose rate of the three reactors. Radiation dose rates decrease to background levels ~3.5 km from the launch site. After a one-year decay time, all of the activated materials produced by an NTR launch would be classified as Class A low-level waste. The activation of air produces significant amounts of argon-41 and nitrogen-16 within 100 m of the launch. The derived air concentration, DAC, from the activation products decays to less than unity within two days, with only argon-41 remaining. After 10 minutes of full power operation the 120 cm core corresponding to a 15 MT payload contains 2.5 x 1013, 1.4 x 1012, 1.5 x 1012, and 7.8 x 10 7 Bq of 131I, 137Cs, 90Sr, and 239Pu respectively. The decay heat after shutdown increases with increasing reactor power with a maximum decay heat of 108 kW immediately after shutdown for the 15 MT payload.

Integrated Studies of Electric Propulsion Engines during Flights in the Earth's Ionosphere

NASA Astrophysics Data System (ADS)

Marov, M. Ya.; Filatyev, A. S.

2018-03-01

Fifty years ago, on October 1, 1966, the first Yantar satellite laboratory with a gas plasma-ion electric propulsion was launched into orbit as part of the Yantar Soviet space program. In 1966-1971, the program launched a total of four laboratories with thrusters operating on argon, nitrogen, and air with jet velocities of 40, 120, and 140 km/s, respectively. These space experiments were the first to demonstrate the long-term stable operation of these thrusters, which exceed chemical rocket engines in specific impulse by an order of magnitude and provide effective jet charge compensation, under the conditions of a real flight at altitudes of 100-400 km. In this article, we have analyzed the potential modern applications of the scientific results obtained by the Yantar space program for the development of air-breathing electric propulsion that ensure the longterm operation of spacecraft in very low orbits.

Megawatt level electric propulsion perspectives

NASA Technical Reports Server (NTRS)

Jahn, Robert G.; Kelly, Arnold J.

1987-01-01

For long range space missions, deliverable payload fraction is an inverse exponential function of the propellant exhaust velocity or specific impulse of the propulsion system. The exhaust velocity of chemical systems are limited by their combustion chemistry and heat transfer to a few km/s. Nuclear rockets may achieve double this range, but are still heat transfer limited and ponderous to develop. Various electric propulsion systems can achieve exhaust velocities in the 10 km/s range, at considerably lower thrust densities, but require an external electrical power source. A general overview is provided of the currently available electric propulsion systems from the perspective of their characteristics as a terminal load for space nuclear systems. A summary of the available electric propulsion options is shown and generally characterized in the power vs. exhaust velocity plot. There are 3 general classes of electric thruster devices: neutral gas heaters, plasma devices, and space charge limited electrostatic or ion thrusters.

ICRF Development for the Variable Specific Impulse Magnetoplasma Rocket

NASA Astrophysics Data System (ADS)

Ryan, P. M.; Baity, F. W.; Barber, G. C.; Carter, M. D.; Hoffman, D. J.; Jaeger, E. F.; Taylor, D. J.; Chang-Diaz, F. R.; Squire, J. P.; McCaskill, G.

1997-11-01

The feasibility of using magnetically vectored and rf-heated plasmas for space propulsion (F. R. Chang-Diaz, et al., Bull. Am. Phys. Soc., 41, 1541 (1996)) is being investigated experimentally on an asymmetric magnetic mirror device at the Advanced Space Propulsion Laboratory (ASPL), Johnson Space Center, NASA. Analysis of the antenna interaction with and the wave propagation through the dense plasma propulsion system is being studied at ORNL(Oak Ridge National Laboratory, managed by Lockheed Martin Energy Research Corp. for the U.S. Department of Energy under contract number DE-AC05-96OR22464.), using antenna design codes developed for ICH systems and mirror codes developed for the EBT experiment at ORNL. The present modeling effort is directed toward the ASPL experimental device. Antenna optimization and performance, as well as the design considerations for space-qualified rf components and systems (minimizing weight while maximizing reliability) will be presented.

Plasma Flow During RF Discharges in VASIMR

NASA Technical Reports Server (NTRS)

Jacobson, V. T.; Chang Diaz, F. R.; Squire, J. P.; Ilin, A. V.; Bengtson, R. D.; Carter, M. D.; Goulding, R. H.

1999-01-01

The Variable Specific Impulse Magnetoplasma Rocket (VASIMR) plasma source consists of a helical antenna, driven at frequencies of 4 to 19 MHz with powers up to 1 kW, in a magnetic field up to 3 kG. Helium is the current test gas, and future experiments with hydrogen are planned. Plasma density and temperature profiles were measured by a reciprocating Langmuir probe, and plasma flow profiles were measured with a reciprocating Mach probe. Both probes were located about 0.5 m downstream from the helical antenna. The plasma source operated in capacitive and inductive modes in addition to a helicon mode. During capacitive and inductive modes, densities were low and plasma flow was < 0.5 Cs. When the plasma operated in a helicon mode, the densities measured downstream from the source were higher [10(exp 12) / cubic cm ] and plasma flow along the magnetic field was of the order Mach 1. Details of the measurements will be shown.

The Benefits of Nuclear Thermal Propulsion (NTP) in an Evolvable Mars Campaign

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Mccurdy, David R.

2014-01-01

NTR: High thrust high specific impulse (2 x LOXLH2chemical) engine uses high power density fission reactor with enriched uranium fuel as thermal power source. Reactor heat is removed using H2propellant which is then exhausted to produce thrust. Conventional chemical engine LH2tanks, turbopumps, regenerative nozzles and radiation-cooled shirt extensions used --NTR is next evolutionary step in high performance liquid rocket engines During the Rover program, a common fuel element tie tube design was developed and used in the design of the 50 klbf Kiwi-B4E (1964), 75 klbf Phoebus-1B (1967), 250 klbf Phoebus-2A (June 1968), then back down to the 25 klbf Pewee engine (Nov-Dec 1968) NASA and DOE are using this same approach: design, build, ground then flight test a small engine using a common fuel element that is scalable to a larger 25 klbf thrust engine needed for human missions

Active space debris removal by a hybrid propulsion module

NASA Astrophysics Data System (ADS)

DeLuca, L. T.; Bernelli, F.; Maggi, F.; Tadini, P.; Pardini, C.; Anselmo, L.; Grassi, M.; Pavarin, D.; Francesconi, A.; Branz, F.; Chiesa, S.; Viola, N.; Bonnal, C.; Trushlyakov, V.; Belokonov, I.

2013-10-01

During the last 40 years, the mass of the artificial objects in orbit increased quite steadily at the rate of about 145 metric tons annually, leading to a total tally of approximately 7000 metric tons. Now, most of the cross-sectional area and mass (97% in LEO) is concentrated in about 4600 intact objects, i.e. abandoned spacecraft and rocket bodies, plus a further 1000 operational spacecraft. Simulations and parametric analyses have shown that the most efficient and effective way to prevent the outbreak of a long-term exponential growth of the catalogued debris population would be to remove enough cross-sectional area and mass from densely populated orbits. In practice, according to the most recent NASA results, the active yearly removal of approximately 0.1% of the abandoned intact objects would be sufficient to stabilize the catalogued debris in low Earth orbit, together with the worldwide adoption of mitigation measures. The candidate targets for removal would have typical masses between 500 and 1000 kg, in the case of spacecraft, and of more than 1000 kg, in the case of rocket upper stages. Current data suggest that optimal active debris removal missions should be carried out in a few critical altitude-inclination bands. This paper deals with the feasibility study of a mission in which the debris is removed by using a hybrid propulsion module as propulsion unit. Specifically, the engine is transferred from a servicing platform to the debris target by a robotic arm so to perform a controlled disposal. Hybrid rocket technology for de-orbiting applications is considered a valuable option due to high specific impulse, intrinsic safety, thrust throttle ability, low environmental impact and reduced operating costs. Typically, in hybrid rockets a gaseous or liquid oxidizer is injected into the combustion chamber along the axial direction to burn a solid fuel. However, the use of tangential injection on a solid grain Pancake Geometry allows for more compact design of the propulsion unit. Only explorative tests were performed in the past on this rocket configuration, which appears to be suitable as de-orbiting system of new satellites as well as for direct application on large debris in the framework of a mission for debris removal. The paper describes some critical aspects of the mission with particular concern to the target selection, the hybrid propulsion module, the operations as well as the systems needed to rendezvous and dock with the target, and the disposal strategy.

Five-Segment Solid Rocket Motor Development Status

NASA Technical Reports Server (NTRS)

Priskos, Alex S.

2012-01-01

In support of the National Aeronautics and Space Administration (NASA), Marshall Space Flight Center (MSFC) is developing a new, more powerful solid rocket motor for space launch applications. To minimize technical risks and development costs, NASA chose to use the Space Shuttle s solid rocket boosters as a starting point in the design and development. The new, five segment motor provides a greater total impulse with improved, more environmentally friendly materials. To meet the mass and trajectory requirements, the motor incorporates substantial design and system upgrades, including new propellant grain geometry with an additional segment, new internal insulation system, and a state-of-the art avionics system. Significant progress has been made in the design, development and testing of the propulsion, and avionics systems. To date, three development motors (one each in 2009, 2010, and 2011) have been successfully static tested by NASA and ATK s Launch Systems Group in Promontory, UT. These development motor tests have validated much of the engineering with substantial data collected, analyzed, and utilized to improve the design. This paper provides an overview of the development progress on the first stage propulsion system.

Development of Advanced Hydrocarbon Fuels at Marshall Space Flight Center

NASA Technical Reports Server (NTRS)

Bai, S. D.; Dumbacher, P.; Cole, J. W.

2002-01-01

This was a small-scale, hot-fire test series to make initial measurements of performance differences of five new liquid fuels relative to rocket propellant-1 (RP-1). The program was part of a high-energy-density materials development at Marshall Space Flight Center (MSFC), and the fuels tested were quadricyclane, 1-7 octodiyne, AFRL-1, biclopropylidene, and competitive impulse noncarcinogenic hypergol (CINCH) (di-methyl-aminoethyl-azide). All tests were conducted at MSFC. The first four fuels were provided by the U.S. Air Force Research Laboratory (AFRL), Edwards Air Force Base, CA. The U.S. Army, Redstone Arsenal, Huntsville, AL, provided the CINCH. The data recorded in all hot-fire tests were used to calculate specific impulse and characteristic exhaust velocity for each fuel, then compared to RP-1 at the same conditions. This was not an exhaustive study, comparing each fuel to RP-1 at an array of mixture ratios, nor did it include important fuel parameters, such as fuel handling or long-term storage. The test hardware was designed for liquid oxygen (lox)/RP-1, then modified for gaseous oxygen/RP-1 to avoid two-phase lox at very small flow rates. All fuels were tested using the same thruster/injector combination designed for RP-1. The results of this test will be used to determine which fuels will be tested in future test programs.

A lightweight pumped hydrazine orbit maneuvering space vehicle

NASA Astrophysics Data System (ADS)

Whitehead, J. C.

1992-01-01

An orbital maneuvering vehicle has a pair of opposed cylindrical piston tanks for hydrazine, and four transverse liquid rocket engines along a longitudinal plane. A new kind of pumped rocket propulsion provides maneuvering thrust on demand, and free-piston pumps which can rapidly start and stop are radially oriented between thrusters. A major advantage of this configuration is that the tanks can be close together, which maximizes the vehicle's longitudinal bending stiffness while minimizing the mass of the central bridging structure. The impulses from pump exhaust and piston reciprocation are directed through the system mass center, so they apply no disturbance torques. All high-temperature components are located on the outside of the central structure, where they are free to expand and radiate heat without detrimental effects. Virtually all lightweight components have been fabricated and tested, and photographs of hardware subassemblies are presented.

Development of Electrothermal Pulsed Plasma Thrusters for Osaka-Institute-of-Technology Electric-Rocket-Engine onboard Small Space Ship

DOE Office of Scientific and Technical Information (OSTI.GOV)

Ishii, Yushuke; Yamamoto, Tsuyoshi; Yamada, Minetsugu

2008-12-31

The Project of Osaka-Institute-of-Technology Electric-Rocket-Engine onboard Small Space Ship (PROITERES) was started at Osaka Institute of Technology. In PROITERES, a 10-kg small satellite with electrothermal pulsed plasma thrusters (PPTs), named JOSHO, will be launched in 2010. The main mission is powered flight of small satellite by electric thruster itself. Electrothermal PPTs were studied with both experiments and numerical simulations. An electrothermal PPT with a side-fed propellant feeding mechanism achieved a total impulse of 3.6 Ns with a repetitive 10000-shot operation. An unsteady numerical simulation showed the existence of considerable amount of ablation delaying to the discharge. However, it was alsomore » shown that this phenomenon should not be regarded as the 'late time ablation' for electrothermal PPTs.« less

Small Launch Vehicle Concept Development for Affordable Multi-Stage Inline Configurations

NASA Technical Reports Server (NTRS)

Beers, Benjamin R.; Waters, Eric D.; Philips, Alan D.; Threet, Grady E. Jr.

2013-01-01

The Advanced Concepts Office at NASA's George C. Marshall Space Flight Center conducted a study of two configurations of a three-stage, inline, liquid propellant small launch vehicle concept developed on the premise of maximizing affordability by targeting a specific payload capability range based on current industry demand. The initial configuration, NESC-1, employed liquid oxygen as the oxidizer and rocket propellant grade kerosene as the fuel in all three stages. The second and more heavily studied configuration, NESC-4, employed liquid oxygen and RP-1 on the first and second stages and liquid oxygen and liquid methane fuel on the third stage. On both vehicles, sensitivity studies were first conducted on specific impulse and stage propellant mass fraction in order to baseline gear ratios and drive the focus of concept development. Subsequent sensitivity and trade studies on the NESC-4 concept investigated potential impacts to affordability due to changes in gross liftoff weight and/or vehicle complexity. Results are discussed at a high level to understand the impact severity of certain sensitivities and how those trade studies conducted can either affect cost, performance, or both.

Trait impulsivity components correlate differently with proactive and reactive control

PubMed Central

Huang, Shihua; Zhu, Zude; Zhang, Wei; Chen, Yu; Zhen, Shuangju

2017-01-01

The relationship between impulsivity and cognitive control is still unknown. We hypothesized that trait impulsivity would differentially correlate with specific cognitive control processes. Trait impulsivity was measured by the Barratt Impulsiveness Scale, which assesses motor, attention, and non-planning impulsiveness components. Cognitive control was measured by a hybrid-designed Stroop task, which distinguishes proactive and reactive control. Thirty-three participants performed the Stroop task while they were scanned by functional magnetic resonance imaging. Proactive and reactive control involved increased activity in the fronto-parietal network, and brain activity was associated with impulsivity scores. Specifically, higher motor impulsiveness was associated with a larger proactive control effect in the inferior parietal lobule and a smaller reactive control effect in the right dorsolateral prefrontal cortex (DLPFC) and anterior cingulate contex. Higher attention impulsivity was associated with a smaller proactive control effect in the right DLPFC. Such a correlation pattern suggests that impulsivity trait components are attributable to different cognitive control subsystems. PMID:28423021

Sex differences in impulsive action and impulsive choice.

PubMed

Weafer, Jessica; de Wit, Harriet

2014-11-01

Here, we review the evidence for sex differences in behavioral measures of impulsivity for both humans and laboratory animals. We focus on two specific components of impulsivity: impulsive action (i.e., difficulty inhibiting a prepotent response) and impulsive choice (i.e., difficulty delaying gratification). Sex differences appear to exist on these measures, but the direction and magnitude of the differences vary. In laboratory animals, impulsive action is typically greater in males than females, whereas impulsive choice is typically greater in females. In humans, women discount more steeply than men, but sex differences on measures of impulsive action depend on tasks and subject samples. We discuss implications of these findings as they relate to drug addiction. We also point out the major gaps in this research to date, including the lack of studies designed specifically to examine sex differences in behavioral impulsivity, and the lack of consideration of menstrual or estrous phase or sex hormone levels in the studies. © 2013.

Sex differences in impulsive action and impulsive choice

PubMed Central

Weafer, Jessica; de Wit, Harriet

2013-01-01

Here, we review the evidence for sex differences in behavioral measures of impulsivity for both humans and laboratory animals. We focus on two specific components of impulsivity: impulsive action (i.e., difficulty inhibiting a prepotent response) and impulsive choice (i.e., difficulty delaying gratification). Sex differences appear to exist on these measures, but the direction and magnitude of the differences vary. In laboratory animals, impulsive action is typically greater in males than females, whereas impulsive choice is typically greater in females. In humans, women discount more steeply than men, but sex differences on measures of impulsive action depend on tasks and subject samples. We discuss implications of these findings as they relate to drug addiction. We also point out the major gaps in this research to date, including the lack of studies designed specifically to examine sex differences in behavioral impulsivity, and the lack of consideration of menstrual or estrous phase or sex hormone levels in the studies. PMID:24286704

Trait impulsivity components correlate differently with proactive and reactive control.

PubMed

Huang, Shihua; Zhu, Zude; Zhang, Wei; Chen, Yu; Zhen, Shuangju

2017-01-01

The relationship between impulsivity and cognitive control is still unknown. We hypothesized that trait impulsivity would differentially correlate with specific cognitive control processes. Trait impulsivity was measured by the Barratt Impulsiveness Scale, which assesses motor, attention, and non-planning impulsiveness components. Cognitive control was measured by a hybrid-designed Stroop task, which distinguishes proactive and reactive control. Thirty-three participants performed the Stroop task while they were scanned by functional magnetic resonance imaging. Proactive and reactive control involved increased activity in the fronto-parietal network, and brain activity was associated with impulsivity scores. Specifically, higher motor impulsiveness was associated with a larger proactive control effect in the inferior parietal lobule and a smaller reactive control effect in the right dorsolateral prefrontal cortex (DLPFC) and anterior cingulate contex. Higher attention impulsivity was associated with a smaller proactive control effect in the right DLPFC. Such a correlation pattern suggests that impulsivity trait components are attributable to different cognitive control subsystems.

Flight motor set 360L001 (STS-26R), volume 1

NASA Technical Reports Server (NTRS)

Ricks, Glen A.

1988-01-01

The NASA space shuttle flight STS-26R, launched at 11:37.00.009 a.m. EDT on 29 Sep. 1988, used the redesigned solid rocket motors (RSRM) 360LOO1A and 360LOO1B. Evaluation of the ground environment instrumentation (GEI) data recorded prior to flight showed: (1) no launch commit criteria violations, (2) that the field joint heater and aft skirt thermal conditioning systems performed adequately, and (3) that the GEI data showed good agreement with thermal model predictions. Evaluation of the developmental flight instrumentation (DFI) revealed excellent agreement with both the predicted and required ballistic specifications. All parameters were well within the CEI specification requirements including propellant burn rates, specific impulse values, and thrust imbalance. Recorded strain values also indicated satisfactory radial growth and stress levels, as well as verification of adequate safety factors. Postflight inspection of the insulation, seals, case, and nozzles showed overall excellent performance. Some thermal DFI protective cork was missing, and inoperative field joint vent valves on the thermal protection cork allowed water entry into the field joints upon splashdown. Evaluation of these anomalies, as well as complete evaluation of all RSRM components, is presented.

The PIT MkV pulsed inductive thruster

NASA Technical Reports Server (NTRS)

Dailey, C. Lee; Lovberg, Ralph H.

1993-01-01

The pulsed inductive thruster (PIT) is an electrodeless, magnetic rocket engine that can operate with any gaseous propellant. A puff of gas injected against the face of a flat (spiral) coil is ionized and ejected by the magnetic field of a fast-rising current pulse from a capacitor bank discharge. Single shot operation on an impulse balance has provided efficiency and I(sub sp) data that characterize operation at any power level (pulse rate). The 1-m diameter MkV thruster concept offers low estimated engine mass at low powers, together with power capability up to more than 1 MW for the 1-m diameter design. A 20 kW design estimate indicates specific mass comparable to Ion Engine specific mass for 10,000 hour operation, while a 100,000 hour design would have a specific mass 1/3 that of the Ion Engine. Performance data are reported for ammonia and hydrazine. With ammonia, at 32 KV coil voltage, efficiency is a little more than 50 percent from 4000 to more than 8000 seconds I(sub sp). Comparison with data at 24 and 28 kV indicates that a wider I(sub sp) range could be achieved at higher coil voltages, if required for deep space missions.

Does impulsivity predict outcome in treatment for binge eating disorder? A multimodal investigation.

PubMed

Manasse, Stephanie M; Espel, Hallie M; Schumacher, Leah M; Kerrigan, Stephanie G; Zhang, Fengqing; Forman, Evan M; Juarascio, Adrienne S

2016-10-01

Multiple dimensions of impulsivity (e.g., affect-driven impulsivity, impulsive inhibition - both general and food-specific, and impulsive decision-making) are associated with binge eating pathology cross-sectionally, yet the literature on whether impulsivity predicts treatment outcome is limited. The present pilot study explored impulsivity-related predictors of 20-week outcome in a small open trial (n = 17) of a novel treatment for binge eating disorder. Overall, dimensions of impulsivity related to emotions (i.e., negative urgency) and food cues emerged as predictors of treatment outcomes (i.e., binge eating frequency and global eating pathology as measured by the Eating Disorders Examination), while more general measures of impulsivity were statistically unrelated to global eating pathology or binge frequency. Specifically, those with higher levels of negative urgency at baseline experienced slower and less pronounced benefit from treatment, and those with higher food-specific impulsivity had more severe global eating pathology at baseline that was consistent at post-treatment and follow-up. These preliminary findings suggest that patients high in negative urgency and with poor response inhibition to food cues may benefit from augmentation of existing treatments to achieve optimal outcomes. Future research will benefit from replication with a larger sample, parsing out the role of different dimensions of impulsivity in treatment outcome for eating disorders, and identifying how treatment can be improved to accommodate higher levels of baseline impulsivity. Copyright © 2016 Elsevier Ltd. All rights reserved.

Electrostatic Plasma Accelerator (EPA)

NASA Technical Reports Server (NTRS)

Brophy, John R.; Aston, Graeme

1995-01-01

The application of electric propulsion to communications satellites, however, has been limited to the use of hydrazine thrusters with electric heaters for thrust and specific impulse augmentation. These electrothermal thrusters operate at specific impulse levels of approximately 300 s with heater powers of about 500 W. Low power arcjets (1-3 kW) are currently being investigated as a way to increase specific impulse levels to approximately 500 s. Ion propulsion systems can easily produce specific impulses of 3000 s or greater, but have yet to be applied to communications satellites. The reasons most often given for not using ion propulsion systems are their high level of overall complexity, low thrust with long burn times, and the difficulty of integrating the propulsion system into existing commercial spacecraft busses. The Electrostatic Plasma Accelerator (EPA) is a thruster concept which promises specific impulse levels between low power arcjets and those of the ion engine while retaining the relative simplicity of the arcjet. The EPA thruster produces thrust through the electrostatic acceleration of a moderately dense plasma. No accelerating electrodes are used and the specific impulse is a direct function of the applied discharge voltage and the propellant atomic mass.

Intrinsic Network Connectivity Patterns Underlying Specific Dimensions of Impulsiveness in Healthy Young Adults.

PubMed

Kubera, Katharina M; Hirjak, Dusan; Wolf, Nadine D; Sambataro, Fabio; Thomann, Philipp A; Wolf, R Christian

2018-05-01

Impulsiveness is a central human personality trait and of high relevance for the development of several mental disorders. Impulsiveness is a multidimensional construct, yet little is known about dimension-specific neural correlates. Here, we address the question whether motor, attentional and non-planning components, as measured by the Barratt Impulsiveness Scale (BIS-11), are associated with distinct or overlapping neural network activity. In this study, we investigated brain activity at rest and its relationship to distinct dimensions of impulsiveness in 30 healthy young adults (m/f = 13/17; age mean/SD = 26.4/2.6 years) using resting-state functional magnetic resonance imaging at 3T. A spatial independent component analysis and a multivariate model selection strategy were used to identify systems loading on distinct impulsivity domains. We first identified eight networks for which we had a-priori hypotheses. These networks included basal ganglia, cortical motor, cingulate and lateral prefrontal systems. From the eight networks, three were associated with impulsiveness measures (p < 0.05, FDR corrected). There were significant relationships between right frontoparietal network function and all three BIS domains. Striatal and midcingulate network activity was associated with motor impulsiveness only. Within the networks regionally confined effects of age and gender were found. These data suggest distinct and overlapping patterns of neural activity underlying specific dimensions of impulsiveness. Motor impulsiveness appears to be specifically related to striatal and midcingulate network activity, in contrast to a domain-unspecific right frontoparietal system. Effects of age and gender have to be considered in young healthy samples.

Predicting the velocity and azimuth of fragments generated by the range destruction or random failure of rocket casings and tankage

NASA Astrophysics Data System (ADS)

Eck, M.; Mukunda, M.

The proliferation of space vehicle launch sites and the projected utilization of these facilities portends an increase in the number of on-pad, ascent, and on-orbit solid-rocket motor (SRM) casings and liquid-rocket tanks which will randomly fail or will fail from range destruct actions. Beyond the obvious safety implications, these failures may have serious resource implications for mission system and facility planners. SRM-casing failures and liquid-rocket tankage failures result in the generation of large, high velocity fragments which may be serious threats to the safety of launch support personnel if proper bunkers and exclusion areas are not provided. In addition, these fragments may be indirect threats to the general public's safety if they encounter hazardous spacecraft payloads which have not been designed to withstand shrapnel of this caliber. They may also become threats to other spacecraft if, by failing on-orbit, they add to the ever increasing space-junk collision cross-section. Most prior attempts to assess the velocity of fragments from failed SRM casings have simply assigned the available chamber impulse to available casing and fuel mass and solved the resulting momentum balance for velocity. This method may predict a fragment velocity which is high or low by a factor of two depending on the ratio of fuel to casing mass extant at the time of failure. Recognizing the limitations of existing methods, the authors devised an analytical approach which properly partitions the available impulse to each major system-mass component. This approach uses the Physics International developed PISCES code to couple the forces generated by an Eulerian modeled gas flow field to a Lagrangian modeled fuel and casing system. The details of a predictive analytical modeling process as well as the development of normalized relations for momentum partition as a function of SRM burn time and initial geometry are discussed in this paper. Methods for applying similar modeling techniques to liquid-tankage-over-pressure failures are also discussed. These methods have been calibrated against observed SRM ascent failures and on-orbit tankage failures. Casing-quadrant sized fragments with velocities exceeding 100 m/s resulted from Titan 34D-SRM range destruct actions at 10 s mission elapsed time (MET). Casing-quadrant sized fragments with velocities of approx. 200 m/s resulted from STS-SRM range destruct actions at 110 s MET. Similar sized fragments for Ariane third stage and Delta second stage tankage were predicted to have maximum velocities of 260 and 480 m/s respectively. Good agreement was found between the predictions and observations for five specific events and it was concluded that the methods developed have good potential for use in predicting the fragmentation process of a number of generically similar casing and tankage systems.

Sleep patterns and impulse control among Japanese junior high school students.

PubMed

Abe, Takeru; Hagihara, Akihito; Nobutomo, Koichi

2010-10-01

Adolescents with decreased impulse control exhibit behavioral problems. Lifestyles are related to impulse control. However, the relations of sleep patterns and impulse control among adolescents are unknown. Thus we examined how sleep patterns were associated with impulse control among Japanese junior high school students. Surveys were completed by a nationwide sample of 1934 students. A significant association between decreased impulse control and bedtimes after midnight was revealed. Specific lifestyle factors related to bedtimes after midnight were older age, greater numbers of hours spent watching television, lack of participation in an extracurricular activity, greater use of convenience stores, and increased attendance at cram schools. This study revealed that going to sleep after midnight was significantly related to decreased impulse control among adolescents. Data about specific lifestyle factors related to going to sleep after midnight should be useful in preventing those behaviors demonstrated by school children that derive from decreased impulse control.

A meta-analytic review of the relationship between adolescent risky sexual behavior and impulsivity across gender, age, and race.

PubMed

Dir, Allyson L; Coskunpinar, Ayca; Cyders, Melissa A

2014-11-01

Impulsivity is frequently included as a risk factor in models of adolescent sexual risk-taking; however, findings on the magnitude of association between impulsivity and risky sexual behavior are variable across studies. The aims of the current meta-analysis were to examine (1) how specific impulsivity traits relate to specific risky sexual behaviors in adolescents, and (2) how the impulsivity-risky sex relationship might differ across gender, age, and race. Eighty-one studies were meta-analyzed using a random effects model to examine the overall impulsivity-risky sex relationship and relationships among specific impulsivity traits and risky sexual behaviors. Overall, results revealed a significant, yet small, association between impulsivity and adolescent risky sexual behavior (r=0.19, p<0.001) that did not differ across impulsivity trait. A pattern of stronger effects was associated with risky sexual behaviors as compared to negative outcomes related to these behaviors. Gender moderated the overall relationship (β=0.22, p=0.04), such that effect sizes were significantly larger in samples with more females. Age, race, study design, and sample type did not moderate the relationship, although there was a pattern suggesting smaller effects for adolescents in juvenile detention settings. Adolescent samples with more females showed a larger impulsivity-risky sex relationship, suggesting that impulsivity may be a more important risk factor for risky sex among adolescent females. Research and treatment should consider gender differences when investigating the role of impulsivity in adolescent sexual risk-taking. Copyright © 2014 Elsevier Ltd. All rights reserved.

Experimental Study of Propulsion Performance by Single-Pulse Rotating Detonation with Gaseous Fuels-Oxygen Mixtures

NASA Astrophysics Data System (ADS)

Toshimitsu, Kazuhiko; Hara, Kosei; Mikajiri, Shuuto; Takiguchi, Naoki

2016-12-01

A rotating detonation engine (RDE) is one of candidates of aerospace engines for supersonic cruse, which is better for propulsion system than a pulse detonation engine (PDE) from the view of continuous thrust and simple structure. The propulsion performance of a proto-type RDE and a PDE by single pulse explosion with methane-oxygen is investigated. Furthermore, the performance of the RDE with acetylene-oxygen gas mixtures is investigated. Its impulse is estimated through ballistic pendulum method with maximum displacement and damping ratio. The comparison of specific impulses of the mixture gases at atmospheric pressure is shown. The specific impulses of the RDE and the PDE are almost same with methane-oxygen gas. Furthermore, the fuel-base specific impulse of the RDE with acetylene-oxygen gas is about over twice as large as one of methane-oxygen, and its maximum specific impulse is 1100 seconds.

Electromagnetic propulsion for spacecraft

NASA Technical Reports Server (NTRS)

Myers, Roger M.

1993-01-01

Three electromagnetic propulsion technologies, solid propellant pulsed plasma thrusters (PPT), magnetoplasmadynamic (MPD) thrusters, and pulsed inductive thrusters (PIT), were developed for application to auxiliary and primary spacecraft propulsion. Both the PPT and MPD thrusters were flown in space, though only PPT's were used on operational satellites. The performance of operational PPT's is quite poor, providing only approximately 8 percent efficiency at approximately 1000 s specific impulse. However, laboratory PPT's yielding 34 percent efficiency at 2000 s specific impulse were extensively tested, and peak performance levels of 53 percent efficiency at 5170 s specific impulse were demonstrated. MPD thrusters were flown as experiments on the Japanese MS-T4 spacecraft and the Space Shuttle and were qualified for a flight in 1994. The flight MPD thrusters were pulsed, with a peak performance of 22 percent efficiency at 2500 s specific impulse using ammonia propellant. Laboratory MPD thrusters were demonstrated with up to 70 percent efficiency and 700 s specific impulse using lithium propellant. While the PIT thruster has never been flown, recent performance measurements using ammonia and hydrazine propellants are extremely encouraging, reaching 50 percent efficiency for specific impulses between 4000 to 8000 s. The fundamental operating principles, performance measurements, and system level design for the three types of electromagnetic thrusters are reviewed, and available data on flight tests are discussed for the PPT and MPD thrusters.

Understanding Impulsivity among Children with Specific Learning Disabilities in Inclusion Schools

ERIC Educational Resources Information Center

Al-Dababneh, Kholoud Adeeb; Al-Zboon, Eman K.

2018-01-01

Impulsive behavior is a characteristic of children with specific learning disabilities (SLD), and is related to learning ability. The present study aims to identify impulsivity behavior in children with SLD who attend inclusion schools, from their resource room teachers' perspectives. A 31-item questionnaire that addressed four subscales was…

Impulsivity, perceived self-regulatory success in dieting, and body mass in children and adolescents: A moderated mediation model.

PubMed

Meule, Adrian; Hofmann, Johannes; Weghuber, Daniel; Blechert, Jens

2016-12-01

Impulsivity has been suggested to contribute to overeating and obesity. However, findings are inconsistent and it appears that only specific facets of impulsivity are related to eating-related variables and to body mass. In the current study, relationships between self-reported impulsivity, perceived self-regulatory success in dieting, and objectively measured body mass were examined in N = 122 children and adolescents. Scores on attentional and motor impulsivity interactively predicted perceived self-regulatory success in dieting, but not body mass: Higher attentional impulsivity was associated with lower perceived self-regulatory success at high levels of motor impulsivity, but not at low levels of motor impulsivity. A moderated mediation model revealed an indirect effect of attentional and motor impulsivity on body mass, which was mediated by perceived self-regulatory success in dieting. Thus, results show that only specific facets of impulsivity are relevant in eating- and weight-regulation and interact with each other in the prediction of these variables. These facets of impulsivity, however, are not directly related to higher body mass, but indirectly via lower success in eating-related self-regulation in children and adolescents. Copyright © 2016 Elsevier Ltd. All rights reserved.

Liquid Oxygen/Liquid Methane Test Summary of the RS-18 Lunar Ascent Engine at Simulated Altitude Conditions at NASA White Sands Test Facility

NASA Technical Reports Server (NTRS)

Melcher, John C., IV; Allred, Jennifer K.

2009-01-01

Tests were conducted with the RS18 rocket engine using liquid oxygen (LO2) and liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA s Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propellant systems in spacecraft applications such as ascent engines or service module engines. Altitude simulation was achieved using the WSTF Large Altitude Simulation System, which provided altitude conditions equivalent up to approx.120,000 ft (approx.37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. Propellant flow rate was measured using a coriolis-style mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup. LO2 flow ranged from 5.9-9.5 lbm/sec (2.7-4.3 kg/sec), and LCH4 flow varied from 3.0-4.4 lbm/sec (1.4-2.0 kg/sec) during the RS-18 hot-fire test series. Thrust was measured using three load cells in parallel. Ignition was demonstrated using a gaseous oxygen/methane spark torch igniter. Data was obtained at multiple chamber pressures, and calculations were performed for specific impulse, C* combustion efficiency, and thrust vector alignment. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/methane lunar ascent engines, 2) obtain vacuum ignition data for the torch and pyrotechnic igniters, and 3) obtain nozzle kinetics data to anchor two-dimensional kinetics codes.

Overlapping and disease specific trait, response, and reflection impulsivity in adolescents with first-episode schizophrenia spectrum disorders or attention-deficit/hyperactivity disorder.

PubMed

Jepsen, J R M; Rydkjaer, J; Fagerlund, B; Pagsberg, A K; Jespersen, R Av F; Glenthøj, B Y; Oranje, B

2018-03-01

Schizophrenia and attention-deficit/hyperactivity disorder (ADHD) are developmental disorders with shared clinical characteristics such as cognitive impairments and impulsivity. Impulsivity is a core feature of ADHD and an important factor in aggression, violence, and substance use in schizophrenia. Based on the hypothesis that schizophrenia and ADHD represent a continuum of neurodevelopmental impairments, the aim was to identify overlapping and disease specific forms of impulsivity. Adolescents between 12 and 17 years of age were assessed with the Schedule for Affective Disorders and Schizophrenia for School-aged Children - Present and Lifetime Version. Subjects with early-onset, first-episode schizophrenia spectrum disorders (EOS) (N = 29) or ADHD (N = 29) and healthy controls (N = 45) were compared on two performance measures (Information Sampling Task, Stop Signal Task) and a subjective personality trait measure of impulsivity (Barratt Impulsiveness Scale, Version 11 (BIS-11)). Significantly increased reflection impulsivity was observed in ADHD but not in the EOS group. No significant response inhibition deficits (stop signal reaction time) were found in the two clinical groups. The ADHD and the EOS group showed significantly increased motor, attentional, and non-planning subtraits of impulsivity. Impaired pre-decisional information gathering appeared to be specific for ADHD while the information gathering was not significantly reduced in subjects with EOS. Neither the ADHD nor EOS group showed impaired response inhibition but shared increased personality subtraits of attentional, non-planning, and motor impulsivity although the latter was significantly more pronounced in ADHD. These increased subtraits of impulsivity may reflect diagnostic non-specific neurodevelopmental impairments in ADHD and EOS in adolescence.

The 260: The Largest Solid Rocket Motor Ever Tested

NASA Technical Reports Server (NTRS)

Crimmins, P.; Cousineau, M.; Rogers, C.; Shell, V.

1999-01-01

Aerojet in the mid 1960s, under contract to NASA, built and static hot fire tested the largest solid rocket motor (SRM) in history for the purpose of demonstrating the feasibility of utilizing large SRMs for space exploration. This program successfully fabricated two high strength steel chambers, loaded each with approximately 1,68 million pounds of propellant, and static test fired these giants with their nozzles up from an underground silo located adjacent to the Florida everglades. Maximum thrust and total impulse in excess of 5,000,000 lbf and 3,470,000,000 lbf-sec were achieved. Flames from the second firing, conducted at night, were seen over eighty miles away. For comparative purposes: the thrust developed was nearly 100 times that of a Minuteman III second stage and the 260 in.-dia cross-section was over 3 times that of the Space Shuttle SRM.

Effect of Stagger on the Vibroacoustic Loads from Clustered Rockets

NASA Technical Reports Server (NTRS)

Rojo, Raymundo; Tinney, Charles E.; Ruf, Joseph H.

2016-01-01

The effect of stagger startup on the vibro-acoustic loads that form during the end- effects-regime of clustered rockets is studied using both full-scale (hot-gas) and laboratory scale (cold gas) data. Both configurations comprise three nozzles with thrust optimized parabolic contours that undergo free shock separated flow and restricted shock separated flow as well as an end-effects regime prior to flowing full. Acoustic pressure waveforms recorded at the base of the nozzle clusters are analyzed using various statistical metrics as well as time-frequency analysis. The findings reveal a significant reduction in end- effects-regime loads when engine ignition is staggered. However, regardless of stagger, both the skewness and kurtosis of the acoustic pressure time derivative elevate to the same levels during the end-effects-regime event thereby demonstrating the intermittence and impulsiveness of the acoustic waveforms that form during engine startup.

Performance Capability of Single-Cavity Vortex Gaseous Nuclear Rockets

NASA Technical Reports Server (NTRS)

Ragsdale, Robert G.

1963-01-01

An analysis was made to determine the maximum powerplant thrust-to-weight ratio possible with a single-cavity vortex gaseous reactor in which all the hydrogen propellant must diffuse through a fuel-rich region. An assumed radial temperature profile was used to represent conduction, convection, and radiation heat-transfer effects. The effect of hydrogen property changes due to dissociation and ionization was taken into account in a hydrodynamic computer program. It is shown that, even for extremely optimistic assumptions of reactor criticality and operating conditions, such a system is limited to reactor thrust-to-weight ratios of about 1.2 x 10(exp -3) for laminar flow. For turbulent flow, the maximum thrust-to-weight ratio is less than 10(exp -3). These low thrusts result from the fact that the hydrogen flow rate is limited by the diffusion process. The performance of a gas-core system with a specific impulse of 3000 seconds and a powerplant thrust-to-weight ratio of 10(exp -2) is shown to be equivalent to that of a 1000-second advanced solid-core system. It is therefore concluded that a single-cavity vortex gaseous reactor in which all the hydrogen must diffuse through the nuclear fuel is a low-thrust device and offers no improvement over a solid-core nuclear-rocket engine. To achieve higher thrust, additional hydrogen flow must be introduced in such a manner that it will by-pass the nuclear fuel. Obviously, such flow must be heated by thermal radiation. An illustrative model of a single-cavity vortex system employing supplementary flow of hydrogen through the core region is briefly examined. Such a system appears capable of thrust-to-weight ratios of approximately 1 to 10. For a high-impulse engine, this capability would be a considerable improvement over solid-core performance. Limits imposed by thermal radiation heat transfer to cavity walls are acknowledged but not evaluated. Alternate vortex concepts that employ many parallel vortices to achieve higher hydrogen flow rates offer the possibility of sufficiently high thrust-to-weight ratios, if they are not limited by short thermal-radiation path lengths.

Nuclear Thermal Rocket (Ntr) Propulsion: A Proven Game-Changing Technology for Future Human Exploration Missions

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; McCurdy, David R.; Packard, Thomas W.

2012-01-01

The NTR represents the next evolutionary step in high performance rocket propulsion. It generates high thrust and has a specific impulse (Isp) of approx.900 seconds (s) or more V twice that of today s best chemical rockets. The technology is also proven. During the previous Rover and NERVA (Nuclear Engine for Rocket Vehicle Applications) nuclear rocket programs, 20 rocket reactors were designed, built and ground tested. These tests demonstrated: (1) a wide range of thrust; (2) high temperature carbide-based nuclear fuel; (3) sustained engine operation; (4) accumulated lifetime; and (5) restart capability V all the requirements needed for a human mission to Mars. Ceramic metal cermet fuel was also pursued, as a backup option. The NTR also has significant growth and evolution potential. Configured as a bimodal system, it can generate electrical power for the spacecraft. Adding an oxygen afterburner nozzle introduces a variable thrust and Isp capability and allows bipropellant operation. In NASA s recent Mars Design Reference Architecture (DRA) 5.0 study, the NTR was selected as the preferred propulsion option because of its proven technology, higher performance, lower launch mass, simple assembly and mission operations. In contrast to other advanced propulsion options, NTP requires no large technology scale-ups. In fact, the smallest engine tested during the Rover program V the 25,000 lbf (25 klbf) Pewee engine is sufficient for human Mars missions when used in a clustered engine arrangement. The Copernicus crewed spacecraft design developed in DRA 5.0 has significant capability and a human exploration strategy is outlined here that uses Copernicus and its key components for precursor near Earth asteroid (NEA) and Mars orbital missions prior to a Mars landing mission. Initially, the basic Copernicus vehicle can enable reusable 1-year round trip human missions to candidate NEAs like 1991 JW and Apophis in the late 2020 s to check out vehicle systems. Afterwards, the Copernicus spacecraft and its 2 key components, now configured as an Earth Return Vehicle / propellant tanker, would be used for a short round trip (approx.18 - 20 months)/short orbital stay (60 days) Mars / Phobos survey mission in 2033 using a split mission approach. The paper also discusses NASA s current Foundational Technology Development activities and its pre-decisional plans for future system-level Technology Demonstrations that include ground testing a small (approx.7.5 klbf) scalable NTR before the decade is out with a flight test shortly thereafter.

Electromagnetic propulsion for spacecraft

NASA Technical Reports Server (NTRS)

Myers, Roger M.

1993-01-01

Three electromagnetic propulsion technologies, solid propellant pulsed plasma thrusters (PPT), magnetoplasmadynamic (MPD) thrusters, and pulsed inductive thrusters (PIT) have been developed for application to auxiliary and primary spacecraft propulsion. Both the PPT and MPD thrusters have been flown in space, though only PPTs have been used on operational satellites. The performance of operational PPTs is quite poor, providing only about 8 percent efficiency at about 1000 sec specific impulse. Laboratory PPTs yielding 34 percent efficiency at 5170 sec specific impulse have been demonstrated. Laboratory MPD thrusters have been demonstrated with up to 70 percent efficiency and 7000 sec specific impulse. Recent PIT performance measurements using ammonia and hydrazine propellants are extremely encouraging, reaching 50 percent efficiency for specific impulses between 4000 and 8000 sec.

Domain-Specific Impulsivity in School-Age Children

PubMed Central

Tsukayama, Eli; Duckworth, Angela Lee; Kim, Betty

2013-01-01

Impulsivity is a salient individual difference in children with well-established predictive validity for life outcomes. The current investigation proposes that impulsive behaviors vary systematically by domain. In a series of studies with ethnically and socioeconomically diverse samples of middle school students, we find that schoolwork-related and interpersonal-related impulsivity, as observed by teachers, parents, and the students themselves, are distinct, moderately correlated behavioral tendencies. Each demonstrates differentiated relationships with dimensions of childhood temperament, Big Five personality factors, and outcomes, such as sociometric popularity, report card grades, and classroom conduct. Implications for theoretical conceptions of impulsivity as well as for practical applications (e.g., domain-specific interventions) are discussed. PMID:24118714

The FOXSI solar sounding rocket campaigns

NASA Astrophysics Data System (ADS)

Glesener, Lindsay; Krucker, Säm.; Christe, Steven; Ishikawa, Shin-nosuke; Buitrago-Casas, Juan Camilo; Ramsey, Brian; Gubarev, Mikhail; Takahashi, Tadayuki; Watanabe, Shin; Takeda, Shin'ichiro; Courtade, Sasha; Turin, Paul; McBride, Stephen; Shourt, Van; Hoberman, Jane; Foster, Natalie; Vievering, Juliana

2016-07-01

The Focusing Optics X-ray Solar Imager (FOXSI) is, in its initial form, a sounding rocket experiment designed to apply the technique of focusing hard X-ray (HXR) optics to the study of fundamental questions about the high-energy Sun. Solar HXRs arise via bremsstrahlung from energetic electrons and hot plasma produced in solar flares and thus are one of the most direct diagnostics of are-accelerated electrons and the impulsive heating of the solar corona. Previous missions have always been limited in sensitivity and dynamic range by the use of indirect (Fourier) imaging due to the lack of availability of direct focusing optics, but technological advances now make direct focusing accessible in the HXR regime (as evidenced by the NuSTAR spacecraft and several suborbital missions). The FOXSI rocket experiment develops and optimizes HXR focusing telescopes for the unique scientific requirements of the Sun. To date, FOXSI has completed two successful flights on 2012 November 02 and 2014 December 11 and is funded for a third flight. This paper gives a brief overview of the experiment, which is sensitive to solar HXRs in the 4-20 keV range, describes its first two flights, and gives a preview of plans for FOXSI-3.

The FOXSI Solar Sounding Rocket Campaigns

NASA Technical Reports Server (NTRS)

Glesener, Lindsay; Krucker, Sam; Christe, Steven; Ishikawa, Shin-Nosuke; Buitrago-Casas, Juan Camilo; Ramsey, Brian; Gubarev, Mikhail; Takahashi, Tadayuki; Watanabe, Shin; Takeda, Shin'ichiro;

The microwave thermal thruster and its application to the launch problem

NASA Astrophysics Data System (ADS)

Parkin, Kevin L. G.

Nuclear thermal thrusters long ago bypassed the 50-year-old specific impulse (Isp) limitation of conventional thrusters, using nuclear powered heat exchangers in place of conventional combustion to heat a hydrogen propellant. These heat exchanger thrusters experimentally achieved an Isp of 825 seconds, but with a thrust-to-weight ratio (T/W) of less than ten they have thus far been too heavy to propel rockets into orbit. This thesis proposes a new idea to achieve both high Isp and high T/W The Microwave Thermal Thruster. This thruster covers the underside of a rocket aeroshell with a lightweight microwave absorbent heat exchange layer that may double as a re-entry heat shield. By illuminating the layer with microwaves directed from a ground-based phased array, an Isp of 700--900 seconds and T/W of 50--150 is possible using a hydrogen propellant. The single propellant simplifies vehicle design, and the high Isp increases payload fraction and structural margins. These factors combined could have a profound effect on the economics of building and reusing rockets. A laboratory-scale microwave thermal heat exchanger is constructed using a single channel in a cylindrical microwave resonant cavity, and new type of coupled electromagnetic-conduction-convection model is developed to simulate it. The resonant cavity approach to small-scale testing reveals several drawbacks, including an unexpected oscillatory behavior. Stable operation of the laboratory-scale thruster is nevertheless successful, and the simulations are consistent with the experimental results. In addition to proposing a new type of propulsion and demonstrating it, this thesis provides three other principal contributions: The first is a new perspective on the launch problem, placing it in a wider economic context. The second is a new type of ascent trajectory that significantly reduces the diameter, and hence cost, of the ground-based phased array. The third is an eclectic collection of data, techniques, and ideas that constitute a Microwave Thermal Rocket as it is presently conceived, in turn selecting and motivating the particular experimental and computational analyses undertaken.

Indirect and direct methods for measuring a dynamic throat diameter in a solid rocket motor

NASA Astrophysics Data System (ADS)

Colbaugh, Lauren

In a solid rocket motor, nozzle throat erosion is dictated by propellant composition, throat material properties, and operating conditions. Throat erosion has a significant effect on motor performance, so it must be accurately characterized to produce a good motor design. In order to correlate throat erosion rate to other parameters, it is first necessary to know what the throat diameter is throughout a motor burn. Thus, an indirect method and a direct method for determining throat diameter in a solid rocket motor are investigated in this thesis. The indirect method looks at the use of pressure and thrust data to solve for throat diameter as a function of time. The indirect method's proof of concept was shown by the good agreement between the ballistics model and the test data from a static motor firing. The ballistics model was within 10% of all measured and calculated performance parameters (e.g. average pressure, specific impulse, maximum thrust, etc.) for tests with throat erosion and within 6% of all measured and calculated performance parameters for tests without throat erosion. The direct method involves the use of x-rays to directly observe a simulated nozzle throat erode in a dynamic environment; this is achieved with a dynamic calibration standard. An image processing algorithm is developed for extracting the diameter dimensions from the x-ray intensity digital images. Static and dynamic tests were conducted. The measured diameter was compared to the known diameter in the calibration standard. All dynamic test results were within +6% / -7% of the actual diameter. Part of the edge detection method consists of dividing the entire x-ray image by an average pixel value, calculated from a set of pixels in the x-ray image. It was found that the accuracy of the edge detection method depends upon the selection of the average pixel value area and subsequently the average pixel value. An average pixel value sensitivity analysis is presented. Both the indirect method and the direct method prove to be viable approaches to determining throat diameter during solid rocket motor operation.

Development and characterization of high-efficiency, high-specific impulse xenon Hall thrusters

NASA Astrophysics Data System (ADS)

Hofer, Richard Robert

This dissertation presents research aimed at extending the efficient operation of 1600 s specific impulse Hall thruster technology to the 2000--3000 s range. While recent studies of commercially developed Hall thrusters demonstrated greater than 4000 s specific impulse, maximum efficiency occurred at less than 3000 s. It was hypothesized that the efficiency maximum resulted as a consequence of modern magnetic field designs, optimized for 1600 s, which were unsuitable at high-specific impulse. Motivated by the industry efforts and mission studies, the aim of this research was to develop and characterize xenon Hall thrusters capable of both high-specific impulse and high-efficiency operation. The research divided into development and characterization phases. During the development phase, the laboratory-model NASA-173M Hall thrusters were designed with plasma lens magnetic field topographies and their performance and plasma characteristics were evaluated. Experiments with the NASA-173M version 1 (v1) validated the plasma lens design by showing how changing the magnetic field topography at high-specific impulse improved efficiency. Experiments with the NASA-173M version 2 (v2) showed there was a minimum current density and optimum magnetic field topography at which efficiency monotonically increased with voltage. Between 300--1000 V, total specific impulse and total efficiency of the NASA-173Mv2 operating at 10 mg/s ranged from 1600--3400 s and 51--61%, respectively. Comparison of the thrusters showed that efficiency can be optimized for specific impulse by varying the plasma lens design. During the characterization phase, additional plasma properties of the NASA-173Mv2 were measured and a performance model was derived accounting for a multiply-charged, partially-ionized plasma. Results from the model based on experimental data showed how efficient operation at high-specific impulse was enabled through regulation of the electron current with the magnetic field. The decrease of efficiency due to multiply-charged ions was minor. Efficiency was largely determined by the current utilization, which suggested maximum Hall thruster efficiency has yet to be reached. The electron Hall parameter was approximately constant with voltage, decreasing from an average of 210 at 300 V to an average of 160 between 400--900 V, which confirmed efficient operation can be realized only over a limited range of Hall parameters.

Impulsivity, Impulsive and Reflective Processes and the Development of Alcohol Use and Misuse in Adolescents and Young Adults

PubMed Central

Wiers, Reinout W.; Ames, Susan L.; Hofmann, Wilhelm; Krank, Marvin; Stacy, Alan W.

2010-01-01

This paper contrasts dual-process and personality approaches in the prediction of addictive behaviors and related risk behaviors. In dual-process models, behavior is described as the joint outcome of qualitatively different “impulsive” (or associative) and “reflective” processes. There are important individual differences regarding both types of processes, and the relative strength of both in a specific situation is influenced by prior behavior and state variables (e.g., fatigue, alcohol use). From this perspective, a specific behavior (e.g., alcohol misuse) can be predicted by the combined indices of the behavior-related impulsive processes (e.g., associations with alcohol), and reflective processes, including the ability to refrain from a motivationally salient action. Personality approaches have reported that general traits such as impulsivity predict addictive behaviors. Here we contrast these two approaches, with supplementary analyses on four datasets. We hypothesized that trait impulsivity can predict specific risky behaviors, but that its predictive power disappears once specific behavior-related associations, indicators of executive functioning, and their interaction are entered into the equation. In all four studies the observed interaction between specific associations and executive control (EC) was robust: trait impulsivity did not diminish the prediction of alcohol use by the interaction. Trait impulsivity was not always related to alcohol use, and when it was, the predictive power disappeared after entering the interaction between behavior-specific associations and EC in one study, but not in the other. These findings are interpreted in relation to the validity of the measurements used, which leads to a more refined hypothesis. PMID:21833213

Dysfunctional, but not functional, impulsivity is associated with a history of seriously violent behaviour and reduced orbitofrontal and hippocampal volumes in schizophrenia.

PubMed

Kumari, Veena; Barkataki, Ian; Goswami, Sangeeta; Flora, Satinder; Das, Mrigendra; Taylor, Pamela

2009-07-15

Aggression and violent acts have been linked with impulsive responding. We investigated whether impulsive personality trait, especially suggestive of dysfunctional impulsivity (i.e. fast and inaccurate responding where this is non-optimal), is associated with a history of seriously violent behaviour and specific brain deficits in schizophrenia. Twenty-four male participants with schizophrenia, of whom 10 had a history of serious physical violence, and 14 healthy male participants were assessed on impulsiveness (dysfunctional impulsivity), venturesomeness (functional impulsivity), and empathy. All participants underwent magnetic resonance imaging. The results revealed that participants with schizophrenia and a history of violence showed elevated impulsiveness but had comparable scores on venturesomeness and empathy dimensions. Impulsiveness scores correlated negatively with reduced orbitofrontal grey matter volume in both the patient and healthy control groups, and with hippocampal volume in the patient group. Our findings suggest that dysfunctional, but not functional, impulsivity is elevated in patients with schizophrenia with a propensity for repetitive violence, and this in turn appears to be associated with reduce volumes of both the orbitofrontal cortex grey matter and the hippocampus. Violence risk prediction and management strategies in schizophrenia may benefit from including specific measures of dysfunctional impulsive traits.

Study of the Helicon Source Operation in the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) Experiment

NASA Technical Reports Server (NTRS)

Molvig, Kim; Batishchev, Oleg

2003-01-01

During this research period the following models of the VASIMR helicon discharge have been further developed and applied to analyze the on-going VX- 10 ASPL experiment: A) 1D semi-analytical model for a mixed-collisional propellant flow B) OD power and balance model for the whole helicon discharge In this particular research period we have concentrated on the MW-level performance of the VASIMR helicon source. Favorable high-power scaling and reduced ionization costs were obtained, and presented at the VASIMR NASA review in the Fall '02. This Grant is continuation of the previous NAG9-1224 award. The research results are summarized in 14 publications; they were presented as 20+ talks at the major International Conferences and scientific seminars at the leading Academic and Research Institutions. The reported results allowed helicon discharge characterization, understanding of the several experimental observations, and helped to make predictions and propose structural modifications for the advanced VASIMR helicon source operation.

SSME lifetime prediction and verification, integrating environments, structures, materials: The challenge

NASA Technical Reports Server (NTRS)

Ryan, R. S.; Salter, L. D.; Young, G. M., III; Munafo, P. M.

1985-01-01

The planned missions for the space shuttle dictated a unique and technology-extending rocket engine. The high specific impulse requirements in conjunction with a 55-mission lifetime, plus volume and weight constraints, produced unique structural design, manufacturing, and verification requirements. Operations from Earth to orbit produce severe dynamic environments, which couple with the extreme pressure and thermal environments associated with the high performance, creating large low cycle loads and high alternating stresses above endurance limit which result in high sensitivity to alternating stresses. Combining all of these effects resulted in the requirements for exotic materials, which are more susceptible to manufacturing problems, and the use of an all-welded structure. The challenge of integrating environments, dynamics, structures, and materials into a verified SSME structure is discussed. The verification program and developmental flight results are included. The first six shuttle flights had engine performance as predicted with no failures. The engine system has met the basic design challenges.

Zero boil-off system testing

NASA Astrophysics Data System (ADS)

Plachta, D. W.; Johnson, W. L.; Feller, J. R.

2016-03-01

Cryogenic propellants such as liquid hydrogen (LH2) and liquid oxygen (LO2) are a part of NASA's future space exploration plans due to their high specific impulse for rocket motors of upper stages. However, the low storage temperatures of LH2 and LO2 cause substantial boil-off losses for long duration missions. These losses can be eliminated by incorporating high performance cryocooler technology to intercept heat load to the propellant tanks and modulating the cryocooler temperature to control tank pressure. The technology being developed by NASA is the reverse turbo-Brayton cycle cryocooler and its integration to the propellant tank through a distributed cooling tubing network coupled to the tank wall. This configuration was recently tested at NASA Glenn Research Center in a vacuum chamber and cryoshroud that simulated the essential thermal aspects of low Earth orbit, its vacuum and temperature. This test series established that the active cooling system integrated with the propellant tank eliminated boil-off and robustly controlled tank pressure.

Advanced space engine preliminary design

NASA Technical Reports Server (NTRS)

Cuffe, J. P. B.; Bradie, R. E.

1973-01-01

A preliminary design was completed for an O2/H2, 89 kN (20,000 lb) thrust staged combustion rocket engine that has a single-bell nozzle with an overall expansion ratio of 400:1. The engine has a best estimate vacuum specific impulse of 4623.8 N-s/kg (471.5 sec) at full thrust and mixture ratio = 6.0. The engine employs gear-driven, low pressure pumps to provide low NPSH capability while individual turbine-driven, high-speed main pumps provide the system pressures required for high-chamber pressure operation. The engine design dry weight for the fixed-nozzle configuration is 206.9 kg (456.3 lb). Engine overall length is 234 cm (92.1 in.). The extendible nozzle version has a stowed length of 141.5 cm (55.7 in.). Critical technology items in the development of the engine were defined. Development program plans and their costs for development, production, operation, and flight support of the ASE were established for minimum cost and minimum time programs.

Gasdynamic Mirror Fusion Propulsion Experiment

NASA Technical Reports Server (NTRS)

Emrich, William J., Jr.; Rodgers, Stephen L. (Technical Monitor)

2001-01-01

Nuclear fusion appears to be the most promising concept for producing extremely high specific impulse rocket engines. One particular fusion concept which seems to be particularly well suited for fusion propulsion applications is the gasdynamic mirror (GDM). This device would operate at much higher plasma densities and with much larger LD ratios than previous mirror machines. Several advantages accrue from such a design. First, the high LA:) ratio minimizes to a large extent certain magnetic curvature effects which lead to plasma instabilities causing a loss of plasma confinement. Second, the high plasma density will result in the plasma behaving much more Re a conventional fluid with a mean free path shorter than the length of the device. This characteristic helps reduce problems associated with "loss cone" microinstabilities. An experimental GDM device is currently being constructed at the NASA Marshall Space Flight Center to provide an initial assessment of the feasibility of this type of propulsion system. Initial experiments are expected to commence in the late fall of 2000.

Zero Boil-Off System Testing

NASA Technical Reports Server (NTRS)

Plachta, D. W.; Johnson, W. L.; Feller, J. R.

2015-01-01

Cryogenic propellants such as liquid hydrogen (LH2) and liquid oxygen (LO2) are a part of NASA's future space exploration plans due to their high specific impulse for rocket motors of upper stages. However, the low storage temperatures of LH2 and LO2 cause substantial boil-off losses for long duration missions. These losses can be eliminated by incorporating high performance cryocooler technology to intercept heat load to the propellant tanks and modulating the cryocooler temperature to control tank pressure. The technology being developed by NASA is the reverse turbo-Brayton cycle cryocooler and its integration to the propellant tank through a distributed cooling tubing network coupled to the tank wall. This configuration was recently tested at NASA Glenn Research Center in a vacuum chamber and cryoshroud that simulated the essential thermal aspects of low Earth orbit, its vacuum and temperature. This test series established that the active cooling system integrated with the propellant tank eliminated boil-off and robustly controlled tank pressure.

Shape Optimization of Supersonic Turbines Using Response Surface and Neural Network Methods

NASA Technical Reports Server (NTRS)

Papila, Nilay; Shyy, Wei; Griffin, Lisa W.; Dorney, Daniel J.

2001-01-01

Turbine performance directly affects engine specific impulse, thrust-to-weight ratio, and cost in a rocket propulsion system. A global optimization framework combining the radial basis neural network (RBNN) and the polynomial-based response surface method (RSM) is constructed for shape optimization of a supersonic turbine. Based on the optimized preliminary design, shape optimization is performed for the first vane and blade of a 2-stage supersonic turbine, involving O(10) design variables. The design of experiment approach is adopted to reduce the data size needed by the optimization task. It is demonstrated that a major merit of the global optimization approach is that it enables one to adaptively revise the design space to perform multiple optimization cycles. This benefit is realized when an optimal design approaches the boundary of a pre-defined design space. Furthermore, by inspecting the influence of each design variable, one can also gain insight into the existence of multiple design choices and select the optimum design based on other factors such as stress and materials considerations.

Space shuttle propulsion estimation development verification

NASA Technical Reports Server (NTRS)

Rogers, Robert M.

1989-01-01

The application of extended Kalman filtering to estimating the Space Shuttle Propulsion performance, i.e., specific impulse, from flight data in a post-flight processing computer program is detailed. The flight data used include inertial platform acceleration, SRB head pressure, SSME chamber pressure and flow rates, and ground based radar tracking data. The key feature in this application is the model used for the SRB's, which is a nominal or reference quasi-static internal ballistics model normalized to the propellant burn depth. Dynamic states of mass overboard and propellant burn depth are included in the filter model to account for real-time deviations from the reference model used. Aerodynamic, plume, wind and main engine uncertainties are also included for an integrated system model. Assuming uncertainty within the propulsion system model and attempts to estimate its deviations represent a new application of parameter estimation for rocket powered vehicles. Illustrations from the results of applying this estimation approach to several missions show good quality propulsion estimates.

Cryogenic Gellant and Fuel Formulation for Metallized Gelled Propellants: Hydrocarbons and Hydrogen with Aluminum

NASA Technical Reports Server (NTRS)

Wong, Wing; Starkovich, John; Adams, Scott; Palaszewski, Bryan; Davison, William; Burt, William; Thridandam, Hareesh; Hu-Peng, Hsiao; Santy, Myrrl J.

1994-01-01

An experimental program to determine the viability of nanoparticulate gellant materials for gelled hydrocarbons and gelled liquid hydrogen was conducted. The gellants included alkoxides (BTMSE and BTMSH) and silica-based materials. Hexane, ethane, propane and hydrogen were gelled with the newly-formulated materials and their rheological properties were determined: shear stress versus shear rate and their attendant viscosities. Metallized hexane with aluminum particles was also rheologically characterized. The propellant and gellant formulations were selected for the very high surface area and relatively-high energy content of the gellants. These new gellants can therefore improve rocket engine specific impulse over that obtained with traditional cryogenic-fuel gellant materials silicon dioxide, frozen methane, or frozen ethane particles. Significant reductions in the total mass of the gellant were enabled in the fuels. In gelled liquid hydrogen, the total mass of gellant was reduced from 10-40 wt percent of frozen hydrocarbon particles to less that 8 wt percent with the alkoxide.

Space rocket engine on the base of the reactor-pumped laser for the interplanetary flights and earth orbital applications

NASA Astrophysics Data System (ADS)

Gulevich, Andrey V.; Dyachenko, Peter P.; Kukharchuk, Oleg F.; Zrodnikov, Anatoly V.

2000-01-01

In this report the concept of vehicle-based reactor-laser engine for long time interplanetary and interorbital (LEO to GEO) flights is proposed. Reactor-pumped lasers offer the perspective way to create on the base of modern nuclear and lasers technologies the low mass and high energy density, repetitively pulsed vehicle-based laser of average power 100 kW. Nowadays the efficiency of nuclear-to-optical energy conversion reached the value of 2-3%. The demo model of reactor-pumped laser facility is under construction in Institute for Physics and Power Engineering (Obninsk, Russia). It enable us to hope that using high power laser on board of the vehicle could make the effective space laser engine possible. Such engine may provide the high specific impulse ~1000-2000 s with the thrust up to 10-100 n. Some calculation results of the characteristics of vehicle-based reactor-laser thermal engine concept are also presented. .

Improved Rhenium Thrust Chambers

NASA Technical Reports Server (NTRS)

O'Dell, John Scott

2015-01-01

Radiation-cooled bipropellant thrust chambers are being considered for ascent/ descent engines and reaction control systems on various NASA missions and spacecraft, such as the Mars Sample Return and Orion Multi-Purpose Crew Vehicle (MPCV). Currently, iridium (Ir)-lined rhenium (Re) combustion chambers are the state of the art for in-space engines. NASA's Advanced Materials Bipropellant Rocket (AMBR) engine, a 150-lbf Ir-Re chamber produced by Plasma Processes and Aerojet Rocketdyne, recently set a hydrazine specific impulse record of 333.5 seconds. To withstand the high loads during terrestrial launch, Re chambers with improved mechanical properties are needed. Recent electrochemical forming (EL-Form"TM") results have shown considerable promise for improving Re's mechanical properties by producing a multilayered deposit composed of a tailored microstructure (i.e., Engineered Re). The Engineered Re processing techniques were optimized, and detailed characterization and mechanical properties tests were performed. The most promising techniques were selected and used to produce an Engineered Re AMBR-sized combustion chamber for testing at Aerojet Rocketdyne.

Muscle-Specific Effective Mechanical Advantage and Joint Impulse in Weightlifting.

PubMed

Kipp, Kristof; Harris, Chad

2017-07-01

Kipp, K, and Harris, C. Muscle-specific effective mechanical advantage and joint impulse in weightlifting. J Strength Cond Res 31(7): 1905-1910, 2017-Lifting greater loads during weightlifting exercises may theoretically be achieved through increasing the magnitudes of net joint impulses or manipulating the joints' effective mechanical advantage (EMA). The purpose of this study was to investigate muscle-specific EMA and joint impulse as well as impulse-momentum characteristics of the lifter-barbell system across a range of external loads during the execution of the clean. Collegiate-level weightlifters performed submaximal cleans at 65, 75, and 85% of their 1-repetition maximum (1-RM), whereas data from a motion analysis system and a force plate were used to calculate lifter-barbell system impulse and velocity, as well as net extensor impulse generated at the hip, knee, and ankle joints and the EMA of the gluteus maximus, hamstrings, quadriceps, and triceps surae muscles. The results indicated that the lifter-barbell system impulse did not change as load increased, whereas the velocity of the lifter-barbell system decreased with greater load. In addition, the net extensor impulse at all joints increased as load increased. The EMA of all muscles did not, however, change as load increased. The load-dependent effects on the impulse-velocity characteristics of the lifter-barbell system may reflect musculoskeletal force-velocity behaviors, and may further indicate that the weightlifting performance is limited by the magnitude of ground reaction force impulse. In turn, the load-dependent effects observed at the joint level indicated that lifting greater loads were due to greater net extensor impulses generated at the joints of the lower extremity and not greater EMAs of the respective extensor muscles. In combination, these results suggest that lifting greater external loads during the clean is due to the ability to generate large extensor joint impulses, rather than manipulate EMA.

A path model of different forms of impulsivity with externalizing and internalizing psychopathology: Towards greater specificity.

PubMed

Johnson, Sheri L; Tharp, Jordan A; Peckham, Andrew D; Carver, Charles S; Haase, Claudia M

2017-09-01

A growing empirical literature indicates that emotion-related impulsivity (compared to impulsivity that is unrelated to emotion) is particularly relevant for understanding a broad range of psychopathologies. Recent work, however, has differentiated two forms of emotion-related impulsivity: A factor termed Pervasive Influence of Feelings captures tendencies for emotions (mostly negative emotions) to quickly shape thoughts, and a factor termed Feelings Trigger Action captures tendencies for positive and negative emotions to quickly and reflexively shape behaviour and speech. This study used path modelling to consider links from emotion-related and non-emotion-related impulsivity to a broad range of psychopathologies. Undergraduates completed self-report measures of impulsivity, depression, anxiety, aggression, and substance use symptoms. A path model (N = 261) indicated specificity of these forms of impulsivity. Pervasive Influence of Feelings was related to anxiety and depression, whereas Feelings Trigger Action and non-emotion-related impulsivity were related to aggression and substance use. The findings of this study suggest that emotion-relevant impulsivity could be a potentially important treatment target for a set of psychopathologies. Recent work has differentiated two forms of emotion-related impulsivity. This study tests a multivariate path model linking emotion-related and non-emotion-related impulsivity with multiple forms of psychopathology. Impulsive thoughts in response to negative emotions were related to anxiety and depression. Impulsive actions in response to emotions were related to aggression and substance use, as did non-emotion-related impulsivity. The study was limited by the reliance on self-report measures of impulsivity and psychopathology. There is a need for longitudinal work on how these forms of impulsivity predict the onset and course of psychopathology. © 2017 The British Psychological Society.

Hypersonic ignition and thrust production in a scramjet

NASA Technical Reports Server (NTRS)

Paull, A.

1993-01-01

Experimental results are given for the specific impulse produced by a two-dimensional scramjet at flight speeds ranging between 2.5 and 5.5 km/s with a combustion chamber Mach number of 4.5. Both hydrogen and ethane fuels were used. Results show that provided sufficiently high pressures and sufficiently long combustion chambers are used specific impulses in excess of 1500 s can be obtained with hydrogen. Ethane produced specific impulses less than 600 s with the same conditions and model configuration.

Performance and Facility Background Pressure Characterization Tests of NASAs 12.5-kW Hall Effect Rocket with Magnetic Shielding Thruster

NASA Technical Reports Server (NTRS)

Kamhawi, Hani; Huang, Wensheng; Haag, Thomas; Shastry, Rohit; Thomas, Robert; Yim, John; Herman, Daniel; Williams, George; Myers, James; Hofer, Richard;

Overweight in adolescent, psychiatric inpatients: A problem of general or food-specific impulsivity?

PubMed

Deux, Natalie; Schlarb, Angelika A; Martin, Franziska; Holtmann, Martin; Hebebrand, Johannes; Legenbauer, Tanja

2017-05-01

Adolescent psychiatric patients are vulnerable to weight problems and show an overrepresentation of overweight compared to the healthy population. One potential factor that can contribute to the etiology of overweight is higher impulsivity. As of yet, it is unclear whether it is a general impulse control deficit or weight-related aspects such as lower impulse control in response to food that have an impact on body weight. As this may have therapeutic implications, the current study investigated differences between overweight and non-overweight adolescent psychiatric inpatients (N = 98; aged 12-20) in relation to trait impulsivity and behavioral inhibition performance. The Barratt Impulsiveness Scale and two go/no-go paradigms with neutral and food-related stimulus materials were applied. Results indicated no significant differences concerning trait impulsivity, but revealed that overweight inpatients had significantly more difficulties in inhibition performance (i.e. they reacted more impulsively) in response to both food and neutral stimuli compared to non-overweight inpatients. Furthermore, no specific inhibition deficit for high-caloric vs. low-caloric food cues emerged in overweight inpatients, whereas non-overweight participants showed significantly lower inhibition skills in response to high-caloric than low-caloric food stimuli. The results highlight a rather general, non-food-specific reduced inhibition performance in an overweight adolescent psychiatric population. Further research is necessary to enhance the understanding of the role of impulsivity in terms of body weight status in this high-risk group of adolescent inpatients. Copyright © 2017 Elsevier Ltd. All rights reserved.

Gender-specific mediational links between parenting styles, parental monitoring, impulsiveness, drinking control, and alcohol-related problems.

PubMed

Patock-Peckham, Julie A; King, Kevin M; Morgan-Lopez, Antonio A; Ulloa, Emilio C; Moses, Jennifer M Filson

2011-03-01

Recently, it has been suggested that traits may dynamically change as conditions change. One possible mechanism that may influence impulsiveness is parental monitoring. Parental monitoring reflects a knowledge regarding one's offspring's whereabouts and social connections. The aim of this investigation was to examine potential gender-specific parental influences to impulsiveness (general behavioral control), control over one's own drinking (specific behavioral control), and alcohol-related problems among individuals in a period of emerging adulthood. Direct and mediational links between parenting styles (permissive, authoritarian, and authoritative), parental monitoring, impulsiveness, drinking control, and alcohol-related problems were investigated. A multiple-group, SEM model with (316 women, 265 men) university students was examined. In general, the overall pattern among male and female respondents was distinct. For daughters, perceptions of a permissive father were indirectly linked to more alcohol-related problems through lower levels of monitoring by fathers and more impulsive symptoms. Perceptions of an authoritative father were also indirectly linked to fewer impulsive symptoms through higher levels of monitoring by fathers among daughters. For men, perceptions of a permissive mother were indirectly linked to more alcohol-related problems through lower levels of monitoring by mothers and more impulsive symptoms. For sons, perceptions of mother authoritativeness were indirectly linked to fewer alcohol-related problems through more monitoring by mothers and fewer impulsive symptoms. Monitoring by an opposite-gender parent mediated the link between parenting styles (i.e., permissive, authoritative) on impulsiveness.

Interactive effects of cumulative stress and impulsivity on alcohol consumption.

PubMed

Fox, Helen C; Bergquist, Keri L; Peihua, Gu; Rajita, Sinha

2010-08-01

Alcohol addiction may reflect adaptations to stress, reward, and regulatory brain systems. While extensive research has identified both stress and impulsivity as independent risk factors for drinking, few studies have assessed the interactive relationship between stress and impulsivity in terms of hazardous drinking within a community sample of regular drinkers. One hundred and thirty regular drinkers (56M/74F) from the local community were assessed for hazardous and harmful patterns of alcohol consumption using the Alcohol Use Disorders Identification Test (AUDIT). All participants were also administered the Barratt Impulsiveness Scale (BIS-11) as a measure of trait impulsivity and the Cumulative Stress/Adversity Checklist (CSC) as a comprehensive measure of cumulative adverse life events. Standard multiple regression models were used to ascertain the independent and interactive nature of both overall stress and impulsivity as well as specific types of stress and impulsivity on hazardous and harmful drinking. Recent life stress, cumulative traumatic stress, overall impulsivity, and nonplanning-related impulsivity as well as cognitive and motor-related impulsivity were all independently predictive of AUDIT scores. However, the interaction between cumulative stress and total impulsivity scores accounted for a significant amount of the variance, indicating that a high to moderate number of adverse events and a high trait impulsivity rating interacted to affect greater AUDIT scores. The subscale of cumulative life trauma accounted for the most variance in AUDIT scores among the stress and impulsivity subscales. Findings highlight the interactive relationship between stress and impulsivity with regard to hazardous drinking. The specific importance of cumulative traumatic stress as a marker for problem drinking is also discussed.

Identifying the facets of impulsivity that explain the relation between ADHD symptoms and substance use in a nonclinical sample

PubMed Central

Roberts, Walter; Peters, Jessica R.; Adams, Zachary W.; Lynam, Donald R.; Milich, Richard

2014-01-01

Adults with attention-deficit/hyperactivity disorder (ADHD) are at higher risk to use substances than their nonclinical peers. Increased levels of impulsivity are generally thought to contribute to their increased levels of risk. Impulsivity is a multifaceted construct, however, and little research to date has attempted to identify which facets of impulsivity contribute to the increased rates of substance abuse among individuals with ADHD. The current study examined the relation among ADHD symptom clusters (i.e., hyperactivity/impulsivity and inattention), substance use rates (i.e., alcohol use, nicotine use, and marijuana use), and personality processes associated with impulsive behavior in a group of young adults. Participants were 361 undergraduate students. Both symptom clusters were positively associated with rates of substance use. Specifically, hyperactive/impulsive symptoms were associated with alcohol and nicotine use, and inattentive symptoms were associated with alcohol use. Several pathways from hyperactive/impulsive symptoms to alcohol, nicotine, and marijuana use via specific facets of impulsivity were identified. These findings have implications for understanding the relation between ADHD symptoms and substance use, as well as clinical implications for preventing and treating substance use problems in individuals with symptoms of ADHD. PMID:24813555

Mission Advantages of Constant Power, Variable Isp Electrostatic Thrusters

NASA Technical Reports Server (NTRS)

Oleson, Steven R.

2000-01-01

Electric propulsion has moved from station-keeping capability for spacecraft to primary propulsion with the advent of both the Deep Space One asteroid flyby and geosynchronous spacecraft orbit insertion. In both cases notably more payload was delivered than would have been possible with chemical propulsion. To provide even greater improvements electrostatic thruster performance could be varied in specific impulse, but kept at constant power to provide better payload or trip time performance for different mission phases. Such variable specific impulse mission applications include geosynchronous and low earth orbit spacecraft stationkeeping and orbit insertion, geosynchronous reusable tug missions, and interplanetary probes. The application of variable specific impulse devices is shown to add from 5 to 15% payload for these missions. The challenges to building such devices include variable voltage power supplies and extending fuel throughput capabilities across the specific impulse range.

Apollo Contour Rocket Nozzle in the Propulsion Systems Laboratory

NASA Image and Video Library

1964-07-21

Bill Harrison and Bud Meilander check the setup of an Apollo Contour rocket nozzle in the Propulsion Systems Laboratory at the National Aeronautics and Space Administration (NASA) Lewis Research Center. The Propulsion Systems Laboratory contained two 14-foot diameter test chambers that could simulate conditions found at very high altitudes. The facility was used in the 1960s to study complex rocket engines such as the Pratt and Whitney RL-10 and rocket components such as the Apollo Contour nozzle, seen here. Meilander oversaw the facility’s mechanics and the installation of test articles into the chambers. Harrison was head of the Supersonic Tunnels Branch in the Test Installations Division. Researchers sought to determine the impulse value of the storable propellant mix, classify and improve the internal engine performance, and compare the results with analytical tools. A special setup was installed in the chamber that included a device to measure the thrust load and a calibration stand. Both cylindrical and conical combustion chambers were examined with the conical large area ratio nozzles. In addition, two contour nozzles were tested, one based on the Apollo Service Propulsion System and the other on the Air Force’s Titan transtage engine. Three types of injectors were investigated, including a Lewis-designed model that produced 98-percent efficiency. It was determined that combustion instability did not affect the nozzle performance. Although much valuable information was obtained during the tests, attempts to improve the engine performance were not successful.

Interactive Effects of Cumulative Stress and Impulsivity on Alcohol Consumption

PubMed Central

Fox, Helen C.; Bergquist, Keri L.; Gu, Peihua; Sinha, Rajita

2013-01-01

Background Alcohol addiction may reflect adaptations to stress, reward, and regulatory brain systems. While extensive research has identified both stress and impulsivity as independent risk factors for drinking, few studies have assessed the interactive relationship between stress and impulsivity in terms of hazardous drinking within a community sample of regular drinkers. Methods One hundred and thirty regular drinkers (56M/74F) from the local community were assessed for hazardous and harmful patterns of alcohol consumption using the Alcohol Use Disorders Identification Test (AUDIT). All participants were also administered the Barratt Impulsiveness Scale (BIS-11) as a measure of trait impulsivity and the Cumulative Stress/Adversity Checklist (CSC) as a comprehensive measure of cumulative adverse life events. Standard multiple regression models were used to ascertain the independent and interactive nature of both overall stress and impulsivity as well as specific types of stress and impulsivity on hazardous and harmful drinking. Results Recent life stress, cumulative traumatic stress, overall impulsivity, and nonplanning-related impulsivity as well as cognitive and motor-related impulsivity were all independently predictive of AUDIT scores. However, the interaction between cumulative stress and total impulsivity scores accounted for a significant amount of the variance, indicating that a high to moderate number of adverse events and a high trait impulsivity rating interacted to affect greater AUDIT scores. The subscale of cumulative life trauma accounted for the most variance in AUDIT scores among the stress and impulsivity subscales. Conclusions Findings highlight the interactive relationship between stress and impulsivity with regard to hazardous drinking. The specific importance of cumulative traumatic stress as a marker for problem drinking is also discussed. PMID:20491738

Attentional impulsivity in binge eating disorder modulates response inhibition performance and frontal brain networks.

PubMed

Hege, M A; Stingl, K T; Kullmann, S; Schag, K; Giel, K E; Zipfel, S; Preissl, H

2015-02-01

A subgroup of overweight and obese people is characterized by binge eating disorder (BED). Increased impulsivity has been suggested to cause binge eating and subsequent weight gain. In the current study, neuronal correlates of increased impulsivity in binge eating disorder during behavioral response inhibition were investigated. Magnetic brain activity and behavioral responses of 37 overweight and obese individuals with and without diagnosed BED were recorded while performing a food-related visual go-nogo task. Trait impulsivity was assessed with the Barratt Impulsiveness Scale (BIS-11). Specifically, increased attentional impulsiveness (a subscale of the BIS-11) in BED was related to decreased response inhibition performance and hypoactivity in the prefrontal control network, which was activated when response inhibition was required. Furthermore, participants with BED showed a trend for a food-specific inhibition performance decline. This was possibly related to the absence of a food-specific activity increase in the prefrontal control network in BED, as observed in the control group. In addition, an increase in activity related to the actual button press during prepotent responses and alterations in visual processing were observed. Our results suggest an attentional impulsiveness-related attenuation in response inhibition performance in individuals with BED. This might have been related to increased reward responsiveness and limited resources to activate the prefrontal control network involved in response inhibition. Our results substantiate the importance of neuronal markers for investigating prevention and treatment of obesity, especially in specific subgroups at risk such as BED.

What Money Can't Buy: Different Patterns in Decision Making About Sex and Money Predict Past Sexual Coercion Perpetration.

PubMed

Carrier Emond, Fannie; Gagnon, Jean; Nolet, Kevin; Cyr, Gaëlle; Rouleau, Joanne-Lucine

2018-02-01

Self-reported impulsivity has been found to predict the perpetration of sexual coercion in both sexual offenders and male college students. Impulsivity can be conceptualized as a generalized lack of self-control (i.e., general perspective) or as a multifaceted construct that can vary from one context to the other (i.e., domain-specific perspective). Delay discounting, the tendency to prefer sooner smaller rewards over larger delayed rewards, is a measure of impulsive decision making. Recent sexual adaptations of delay discounting tasks can be used to test domain-specific assumptions. The present study used the UPPS-P impulsivity questionnaire, a standard money discounting task, and a sexual discounting task to predict past use of sexual coercion in a sample of 98 male college students. Results indicated that higher negative urgency scores, less impulsive money discounting, and more impulsive sexual discounting all predicted sexual coercion. Consistent with previous studies, sexuality was discounted more steeply than money by both perpetrators and non-perpetrators of sexual coercion, but this difference was twice as large in perpetrators compared to non-perpetrators. Our study identified three different predictors of sexual coercion in male college students: a broad tendency to act rashly under negative emotions, a specific difficulty to postpone sexual gratification, and a pattern of optimal non-sexual decision making. Results highlight the importance of using multiple measures, including sexuality-specific measures, to get a clear portrait of the links between impulsivity and sexual coercion.

Enrichment Zoning Options for the Small Nuclear Rocket Engine (SNRE)

DOE Office of Scientific and Technical Information (OSTI.GOV)

Bruce G. Schnitzler; Stanley K. Borowski

2010-07-01

Advancement of U.S. scientific, security, and economic interests through a robust space exploration program requires high performance propulsion systems to support a variety of robotic and crewed missions beyond low Earth orbit. In NASA’s recent Mars Design Reference Architecture (DRA) 5.0 study (NASA-SP-2009-566, July 2009), nuclear thermal propulsion (NTP) was again selected over chemical propulsion as the preferred in-space transportation system option because of its high thrust and high specific impulse (-900 s) capability, increased tolerance to payload mass growth and architecture changes, and lower total initial mass in low Earth orbit. An extensive nuclear thermal rocket technology development effortmore » was conducted from 1955-1973 under the Rover/NERVA Program. The Small Nuclear Rocket Engine (SNRE) was the last engine design studied by the Los Alamos National Laboratory during the program. At the time, this engine was a state-of-the-art design incorporating lessons learned from the very successful technology development program. Past activities at the NASA Glenn Research Center have included development of highly detailed MCNP Monte Carlo transport models of the SNRE and other small engine designs. Preliminary core configurations typically employ fuel elements with fixed fuel composition and fissile material enrichment. Uniform fuel loadings result in undesirable radial power and temperature profiles in the engines. Engine performance can be improved by some combination of propellant flow control at the fuel element level and by varying the fuel composition. Enrichment zoning at the fuel element level with lower enrichments in the higher power elements at the core center and on the core periphery is particularly effective. Power flattening by enrichment zoning typically results in more uniform propellant exit temperatures and improved engine performance. For the SNRE, element enrichment zoning provided very flat radial power profiles with 551 of the 564 fuel elements within 1% of the average element power. Results for this and alternate enrichment zoning options for the SNRE are compared.« less

A General Approach to the Geostationary Transfer Orbit Mission Recovery

NASA Technical Reports Server (NTRS)

Faber, Nicolas; Aresini, Andrea; Wauthier, Pascal; Francken, Philippe

2007-01-01

This paper discusses recovery scenarios for geosynchronous satellites injected in a non-nominal orbit due to a launcher underperformance. The theory on minimum-fuel orbital transfers is applied to develop an operational tool capable to design a recovery mission. To obtain promising initial guesses for the recovery three complementary techniques are used: p-optimized impulse function contouring, a numerical impulse function minimization and the solutions to the switching equations. The tool evaluates the feasibility of a recovery with the on-board propellant of the spacecraft and performs the complete mission design. This design takes into account for various mission operational constraints such as e.g., the requirement of multiple finite-duration burns, third-body orbital perturbations, spacecraft attitude constraints and ground station visibility. In a final case study, we analyze the consequences of a premature breakdown of an upper rocket stage engine during injection on a geostationary transfer orbit, as well as the possible recovery solution with the satellite on-board propellant.

Long life monopropellant hydrazine thruster evaluation for Space Station Freedom application - Test results

NASA Technical Reports Server (NTRS)

Popp, Christopher G.; Cook, Joseph C.; Ragland, Brenda L.; Pate, Leah R.

1992-01-01

In support of propulsion system thruster development activity for Space Station Freedom (SSF), NASA Johnson Space Center (JSC) conducted a hydrazine thruster technology demonstration program. The goal of this program was to identify impulse life capability of state-of-the-art long life hydrazine thrusters nominally rated for 50 pounds thrust at 300 psia supply pressure. The SSF propulsion system requirement for impulse life of this thruster class is 1.5 million pounds-seconds, corresponding to a throughput of approximately 6400 pounds of propellant. Long life thrusters were procured from The Marquardt Company, Hamilton Standard, and Rocket Research Company, Testing at JSC was completed on the thruster designs to quantify life while simulating expected thruster firing duty cycles and durations for SSF. This paper presents a review of the SSF propulsion system hydrazine thruster requirements, summaries of the three long life thruster designs procured by JSC and acceptance test results for each thruster, the JSC thruster life evaluation test program, and the results of the JSC test program.

Results of a space shuttle pulme impingement investigation at stage separation in the NASA-MSFC impulse base flow facility

NASA Technical Reports Server (NTRS)

Mccanna, R. W.; Sims, W. H.

1972-01-01

Results are presented for an experimental space shuttle stage separation plume impingement program conducted in the NASA-Marshall Space Flight Center's impulse base flow facility (IBFF). Major objectives of the investigation were to: (1)determine the degree of dual engine exhaust plume simulation obtained using the equivalent engine; (2) determine the applicability of the analytical techniques; and (3) obtain data applicable for use in full-scale studies. The IBFF tests determined the orbiter rocket motor plume impingement loads, both pressure and heating, on a 3 percent General Dynamics B-15B booster configuration in a quiescent environment simulating a nominal staging altitude of 73.2 km (240,00 ft). The data included plume surveys of two 3 percent scale orbiter nozzles, and a 4.242 percent scaled equivalent nozzle - equivalent in the sense that it was designed to have the same nozzle-throat-to-area ratio as the two 3 percent nozzles and, within the tolerances assigned for machining the hardware, this was accomplished.

Applied-field MPD thruster geometry effects

NASA Technical Reports Server (NTRS)

Myers, Roger M.

1991-01-01

Eight MPD thruster configurations were used to study the effects of applied field strength, propellant, and facility pressure on thruster performance. Vacuum facility background pressures higher than approx. 0.12 Pa were found to greatly influence thruster performance and electrode power deposition. Thrust efficiency and specific impulse increased monotonically with increasing applied field strength. Both cathode and anode radii fundamentally influenced the efficiency specific impulse relationship, while their lengths influence only the magnitude of the applied magnetic field required to reach a given performance level. At a given specific impulse, large electrode radii result in lower efficiencies for the operating conditions studied. For all test conditions, anode power deposition was the largest efficiency loss, and represented between 50 and 80 pct. of the input power. The fraction of the input power deposited into the anode decreased with increasing applied field and anode radii. The highest performance measured, 20 pct. efficiency at 3700 seconds specific impulse, was obtained using hydrogen propellant.

Angular distributions of plasma edge velocity and integrated intensity: Update on specific impulse for Ablative Laser Propulsion

NASA Astrophysics Data System (ADS)

Lin, Jun; Pakhomov, Andrew V.

2005-04-01

This work concludes our discussion of the image processing technique developed earlier for determination of specific impulse (Isp) for Ablative Laser Propulsion (ALP). The plasma plumes are recorded with a time-resolved intensified charge-coupled device (ICCD) camera. The plasma was formed in vacuum (˜ 3×10-3 Torr) by focusing output pulses of a laser system (100-ps pulsewidth at 532 nm wavelength and ˜35 mJ energy) on surfaces of C (graphite), Al, Si, Fe, Cu, Zn, Sn, and Pb elements. Angular profiles for integrated intensity and plasma expansion velocity were determined for the tested elements. Such profiles were used further for assessment of specific impulse. Specific impulses derived from angular distributions of plasma expansion velocity and integral intensity appeared in excellent agreement with the data derived earlier from force measurements.

Choice Impulsivity: Definitions, Measurement Issues, and Clinical Implications

PubMed Central

Hamilton, Kristen R.; Mitchell, Marci R.; Wing, Victoria C.; Balodis, Iris M.; Bickel, Warren K.; Fillmore, Mark; Lane, Scott D.; Lejuez, C. W.; Littlefield, Andrew K.; Luijten, Maartje; Mathias, Charles W.; Mitchell, Suzanne H.; Napier, T. Celeste; Reynolds, Brady; Schütz, Christian G.; Setlow, Barry; Sher, Kenneth J.; Swann, Alan C.; Tedford, Stephanie E.; White, Melanie J.; Winstanley, Catharine A.; Yi, Richard; Potenza, Marc N.; Moeller, F. Gerard

2015-01-01

Background Impulsivity critically relates to many psychiatric disorders. Given the multi-faceted construct that impulsivity represents, defining core aspects of impulsivity is vital for the assessment and understanding of clinical conditions. Choice impulsivity (CI), involving the preferential selection of smaller sooner rewards over larger later rewards, represents one important type of impulsivity. Method The International Society for Research on Impulsivity (InSRI) convened to discuss the definition and assessment of CI and provide recommendations regarding measurement across species. Results Commonly used preclinical and clinical CI behavioral tasks are described, and considerations for each task are provided to guide CI task selection. Differences in assessment of CI (self-report, behavioral) and calculating CI indices (e.g., area-under-the-curve, indifference point, steepness of discounting curve) are discussed along with properties of specific behavioral tasks used in preclinical and clinical settings. Conclusions The InSRI group recommends inclusion of measures of CI in human studies examining impulsivity. Animal studies examining impulsivity should also include assessments of CI and these measures should be harmonized in accordance with human studies of the disorders being modeled in the preclinical investigations. The choice of specific CI measures to be included should be based on the goals of the study and existing preclinical and clinical literature using established CI measures. PMID:25867841

Impulsive suicide attempts: a systematic literature review of definitions, characteristics and risk factors.

PubMed

Rimkeviciene, Jurgita; O'Gorman, John; De Leo, Diego

2015-01-15

Extensive research on impulsive suicide attempts, but lack of agreement on the use of this term indicates the need for a systematic literature review of the area. The aim of this review was to examine definitions and likely correlates of impulsive attempts. A search of Medline, Psychinfo, Scopus, Proquest and Web of Knowledge databases was conducted. Additional articles were identified using the cross-referencing function of Google Scholar. 179 relevant papers were identified. Four different groups of research criteria used to assess suicide attempt impulsivity emerged: (a) time-related criteria, (b) absence of proximal planning/preparations, (c) presence of suicide plan in lifetime/previous year, and (d) other. Subsequent analysis used these criteria to compare results from different studies on 20 most researched hypotheses. Conclusions regarding the characteristics of impulsive attempts are more consistent than those on the risk factors specific to such attempts. No risk factors were identified that uniformly related to suicide attempt impulsivity across all criteria groups, but relationships emerged between separate criteria and specific characteristics of suicide attempters. Only published articles were included. Large inconsistencies in methods of the studies included in this review prevented comparison of effect sizes. The vast disparities in findings on risk factors for impulsive suicide attempts among different criteria groups suggest the need to address the methodological issues in defining suicide attempt impulsivity before further research into correlates of such attempts can effectively progress. Specific recommendations are offered for necessary research. Copyright © 2014 Elsevier B.V. All rights reserved.

Advanced Power and Propulsion: Insuring Human Survival and Productivity in Deep Space Missions

NASA Technical Reports Server (NTRS)

Chang-Diaz, Franklin R.

2001-01-01

Dr. Chang-Diaz gave an intriguing presentation of his research in advanced rocket propulsion and its relevance for planning and executing crewed deep space explorations. Though not necessarily exclusively Martian, his thrust looks critically at future Mars missions. Initially Dr. Chang-Diaz showed the time constraints of Mars missions due to orbital mechanics and our present chemically powered rocket technology. Since essentially all the energy required to place current generation spacecraft into a Martian trajectory must be expended in the early minutes of a flight, most of such a mission is spent in free-fall drift, captive to the gravitational forces among Earth, the Sun, and Mars. The simple physics of such chemically powered missions requires nearly a year in transit for each direction of a Mars mission. And the optimal orientations of Earth and Mars for rendezvous require further time on or around Mars to await return. These extensions of mission duration place any crew under a three-fold jeopardy: (1) physiological deconditioning (which in some aspects is still unknown and unpreventable), (2) psychological stress, and (3) ionizing radiation. This latter risk is due to exposure of crew members for extended time to the highly unpredictable and potentially lethal radiations of open space. Any gains in shortening mission duration would reap equivalent or greater benefits for these crew concerns. Dr. Chang-Diaz has applied his training and expertise (Ph.D. from Massachusetts Institute of Technology in applied plasma physics) toward development of continuous rocket propulsion which would offer great time advantages in travel, and also more launch options than are now available. He clearly explained the enormous gains from a relatively low thrust accelerative force applied essentially continuously versus the high, but short-lived propulsion of present chemical rockets. In fact, such spacecraft could be powered throughout the mission, accelerating to approximately the midpoint and decelerating during the latter half. This would not only provide some level of gravity (acceleration) throughout the mission but also allow very high velocities to be achieved, thus saving many months of travel time. In proposing the design of such a spacecraft propulsion system, Dr. Chang-Diaz was quick to acknowledge the need for a large power source, which undoubtedly must be nuclear fueled at the solar distances involved. He calls his system the Variable Specific Impulse Magnetoplasma Rocket (VASIMR). The other major ingredient is a mass (deuterium, which could also function as a radiation shield for crews) for energizing into the ultra hot, high velocity exhaust plasma. He foresees models now functional in the laboratory soon to be tested in space. In fact, some of these concepts have already been tried there. His optimism and determination would have operational rockets in the next decades.

Comparison between model predictions and observations of ELF radio atmospherics generated by rocket-triggered lightning

NASA Astrophysics Data System (ADS)

Dupree, N. A.; Moore, R. C.

2011-12-01

Model predictions of the ELF radio atmospheric generated by rocket-triggered lightning are compared with observations performed at Arrival Heights, Antarctica. The ability to infer source characteristics using observations at great distances may prove to greatly enhance the understanding of lightning processes that are associated with the production of transient luminous events (TLEs) as well as other ionospheric effects associated with lightning. The modeling of the sferic waveform is carried out using a modified version of the Long Wavelength Propagation Capability (LWPC) code developed by the Naval Ocean Systems Center over a period of many years. LWPC is an inherently narrowband propagation code that has been modified to predict the broadband response of the Earth-ionosphere waveguide to an impulsive lightning flash while preserving the ability of LWPC to account for an inhomogeneous waveguide. ELF observations performed at Arrival Heights, Antarctica during rocket-triggered lightning experiments at the International Center for Lightning Research and Testing (ICLRT) located at Camp Blanding, Florida are presented. The lightning current waveforms directly measured at the base of the lightning channel (at the ICLRT) are used together with LWPC to predict the sferic waveform observed at Arrival Heights under various ionospheric conditions. This paper critically compares observations with model predictions.

Engine Cycle Analysis of Air Breathing Microwave Rocket with Reed Valves

DOE Office of Scientific and Technical Information (OSTI.GOV)

Fukunari, Masafumi; Komatsu, Reiji; Yamaguchi, Toshikazu

The Microwave Rocket is a candidate for a low cost launcher system. Pulsed plasma generated by a high power millimeter wave beam drives a blast wave, and a vehicle acquires impulsive thrust by exhausting the blast wave. The thrust generation process of the Microwave Rocket is similar to a pulse detonation engine. In order to enhance the performance of its air refreshment, the air-breathing mechanism using reed valves is under development. Ambient air is taken to the thruster through reed valves. Reed valves are closed while the inside pressure is high enough. After the time when the shock wave exhaustsmore » at the open end, an expansion wave is driven and propagates to the thrust-wall. The reed valve is opened by the negative gauge pressure induced by the expansion wave and its reflection wave. In these processes, the pressure oscillation is important parameter. In this paper, the pressure oscillation in the thruster was calculated by CFD combined with the flux through from reed valves, which is estimated analytically. As a result, the air-breathing performance is evaluated using Partial Filling Rate (PFR), the ratio of thruster length to diameter L/D, and ratio of opening area of reed valves to superficial area {alpha}. An engine cycle and predicted thrust was explained.« less

Effects of cannabis on impulsivity: a systematic review of neuroimaging findings.

PubMed

Wrege, Johannes; Schmidt, Andre; Walter, Anna; Smieskova, Renata; Bendfeldt, Kerstin; Radue, Ernst-Wilhelm; Lang, Undine E; Borgwardt, Stefan

2014-01-01

We conducted a systematic review to assess the evidence for specific effects of cannabis on impulsivity, disinhibition and motor control. The review had a specific focus on neuroimaging findings associated with acute and chronic use of the drug and covers literature published up until May 2012. Seventeen studies were identified, of which 13 met the inclusion criteria; three studies investigated acute effects of cannabis (1 fMRI, 2 PET), while six studies investigated non-acute functional effects (4 fMRI, 2 PET), and four studies investigated structural alterations. Functional imaging studies of impulsivity studies suggest that prefrontal blood flow is lower in chronic cannabis users than in controls. Studies of acute administration of THC or marijuana report increased brain metabolism in several brain regions during impulsivity tasks. Structural imaging studies of cannabis users found differences in reduced prefrontal volumes and white matter integrity that might mediate the abnormal impulsivity and mood observed in marijuana users. To address the question whether impulsivity as a trait precedes cannabis consumption or whether cannabis aggravates impulsivity and discontinuation of usage more longitudinal study designs are warranted.

Effects of Cannabis on Impulsivity: A Systematic Review of Neuroimaging Findings

PubMed Central

Wrege, Johannes; Schmidt, André; Walter, Anna; Smieskova, Renata; Bendfeldt, Kerstin; Radue, Ernst-Wilhelm; Lang, Undine E.; Borgwardt, Stefan

2014-01-01

We conducted a systematic review to assess the evidence for specific effects of cannabis on impulsivity, disinhibition and motor control. The review had a specific focus on neuroimaging findings associated with acute and chronic use of the drug and covers literature published up until May 2012. Seventeen studies were identified, of which 13 met the inclusion criteria; three studies investigated acute effects of cannabis (1 fMRI, 2 PET), while six studies investigated non-acute functional effects (4 fMRI, 2 PET), and four studies investigated structural alterations. Functional imaging studies of impulsivity studies suggest that prefrontal blood flow is lower in chronic cannabis users than in controls. Studies of acute administration of THC or marijuana report increased brain metabolism in several brain regions during impulsivity tasks. Structural imaging studies of cannabis users found differences in reduced prefrontal volumes and white matter integrity that might mediate the abnormal impulsivity and mood observed in marijuana users. To address the question whether impulsivity as a trait precedes cannabis consumption or whether cannabis aggravates impulsivity and discontinuation of usage more longitudinal study designs are warranted. PMID:23829358

Development and Characterization of High-Efficiency, High-Specific Impulse Xenon Hall Thrusters

NASA Technical Reports Server (NTRS)

Hofer, Richard R.; Jacobson, David (Technical Monitor)

2004-01-01

This dissertation presents research aimed at extending the efficient operation of 1600 s specific impulse Hall thruster technology to the 2000 to 3000 s range. Motivated by previous industry efforts and mission studies, the aim of this research was to develop and characterize xenon Hall thrusters capable of both high-specific impulse and high-efficiency operation. During the development phase, the laboratory-model NASA 173M Hall thrusters were designed and their performance and plasma characteristics were evaluated. Experiments with the NASA-173M version 1 (v1) validated the plasma lens magnetic field design. Experiments with the NASA 173M version 2 (v2) showed there was a minimum current density and optimum magnetic field topography at which efficiency monotonically increased with voltage. Comparison of the thrusters showed that efficiency can be optimized for specific impulse by varying the plasma lens. During the characterization phase, additional plasma properties of the NASA 173Mv2 were measured and a performance model was derived. Results from the model and experimental data showed how efficient operation at high-specific impulse was enabled through regulation of the electron current with the magnetic field. The electron Hall parameter was approximately constant with voltage, which confirmed efficient operation can be realized only over a limited range of Hall parameters.

A centre-triggered magnesium fuelled cathodic arc thruster uses sublimation to deliver a record high specific impulse

NASA Astrophysics Data System (ADS)

Neumann, Patrick R. C.; Bilek, Marcela; McKenzie, David R.

2016-08-01

The cathodic arc is a high current, low voltage discharge that operates in vacuum and provides a stream of highly ionised plasma from a solid conducting cathode. The high ion velocities, together with the high ionisation fraction and the quasineutrality of the exhaust stream, make the cathodic arc an attractive plasma source for spacecraft propulsion applications. The specific impulse of the cathodic arc thruster is substantially increased when the emission of neutral species is reduced. Here, we demonstrate a reduction of neutral emission by exploiting sublimation in cathode spots and enhanced ionisation of the plasma in short, high-current pulses. This, combined with the enhanced directionality due to the efficient erosion profiles created by centre-triggering, substantially increases the specific impulse. We present experimentally measured specific impulses and jet power efficiencies for titanium and magnesium fuels. Our Mg fuelled source provides the highest reported specific impulse for a gridless ion thruster and is competitive with all flight rated ion thrusters. We present a model based on cathode sublimation and melting at the cathodic arc spot explaining the outstanding performance of the Mg fuelled source. A further significant advantage of an Mg-fuelled thruster is the abundance of Mg in asteroidal material and in space junk, providing an opportunity for utilising these resources in space.

Correlation of rocket propulsion fuel properties with chemical composition using comprehensive two-dimensional gas chromatography with time-of-flight mass spectrometry followed by partial least squares regression analysis.

PubMed

Kehimkar, Benjamin; Hoggard, Jamin C; Marney, Luke C; Billingsley, Matthew C; Fraga, Carlos G; Bruno, Thomas J; Synovec, Robert E

2014-01-31

There is an increased need to more fully assess and control the composition of kerosene-based rocket propulsion fuels such as RP-1. In particular, it is critical to make better quantitative connections among the following three attributes: fuel performance (thermal stability, sooting propensity, engine specific impulse, etc.), fuel properties (such as flash point, density, kinematic viscosity, net heat of combustion, and hydrogen content), and the chemical composition of a given fuel, i.e., amounts of specific chemical compounds and compound classes present in a fuel as a result of feedstock blending and/or processing. Recent efforts in predicting fuel chemical and physical behavior through modeling put greater emphasis on attaining detailed and accurate fuel properties and fuel composition information. Often, one-dimensional gas chromatography (GC) combined with mass spectrometry (MS) is employed to provide chemical composition information. Building on approaches that used GC-MS, but to glean substantially more chemical information from these complex fuels, we recently studied the use of comprehensive two dimensional (2D) gas chromatography combined with time-of-flight mass spectrometry (GC×GC-TOFMS) using a "reversed column" format: RTX-wax column for the first dimension, and a RTX-1 column for the second dimension. In this report, by applying chemometric data analysis, specifically partial least-squares (PLS) regression analysis, we are able to readily model (and correlate) the chemical compositional information provided by use of GC×GC-TOFMS to RP-1 fuel property information such as density, kinematic viscosity, net heat of combustion, and so on. Furthermore, we readily identified compounds that contribute significantly to measured differences in fuel properties based on results from the PLS models. We anticipate this new chemical analysis strategy will have broad implications for the development of high fidelity composition-property models, leading to an improved approach to fuel formulation and specification for advanced engine cycles. Copyright © 2014 Elsevier B.V. All rights reserved.

Active seismic experiment

NASA Technical Reports Server (NTRS)

Kovach, R. L.; Watkins, J. S.; Talwani, P.

1972-01-01

The Apollo 16 active seismic experiment (ASE) was designed to generate and monitor seismic waves for the study of the lunar near-surface structure. Several seismic energy sources are used: an astronaut-activated thumper device, a mortar package that contains rocket-launched grenades, and the impulse produced by the lunar module ascent. Analysis of some seismic signals recorded by the ASE has provided data concerning the near-surface structure at the Descartes landing site. Two compressional seismic velocities have so far been recognized in the seismic data. The deployment of the ASE is described, and the significant results obtained are discussed.

The impact of self-reported life stress on current impulsivity in cocaine dependent adults

PubMed Central

Ross, Elizabeth L.; Yoon, Jin H.; Mahoney, James J.; Omar, Yasmine; Newton, Thomas F.; De La Garza, Richard

2014-01-01

Current cocaine treatments may be enhanced with a better understanding of the underlying mechanisms that contribute to the onset and maintenance of the disease, such as life stress and impulsivity. Life stress and impulsivity have previously been studied independently as contributors to drug use, and the current study expands upon past research by examining how these factors interact with one another. The aim of the current study was to evaluate the role of life stress in predicting impulsivity in a non-treatment seeking cocaine-dependent sample (N = 112). Analyses revealed that trait impulsivity (as measured by the Barratt Impulsiveness Scale) was associated with education (r = −3.09, p < 0.01), as those who had higher educational attainment also reported lower rates of trait impulsivity. In addition, those over the age of 30 demonstrated lower impulsivity in decision-making (as measured by delay discounting) than those under 30 (t = 2.21, p = 0.03). Overall exposure to life stress was not significantly correlated to either aspect of impulsivity. However several specific life stressors were significantly related to greater impulsivity including having been put up for adoption or in foster care (t = −2.96, p < 0.01), and having a child taken away against their will (t = −2.68, p = 0.01). These findings suggest that age and education relate to impulsivity; and that while an overall compilation of life stress scores was not related to impulsivity, specific types of stress related to either being taken away from a parent or having a child taken away were. Future studies should assess these constructs longitudinally to restrict response bias. PMID:23796525

Development of a 12-Thrust Chamber Kerosene /Oxygen Primary Rocket Sub-System for an Early (1964) Air-Augmented Rocket Ground-Test System

NASA Technical Reports Server (NTRS)

Pryor, D.; Hyde, E. H.; Escher, W. J. D.

1999-01-01

Airbreathing/Rocket combined-cycle, and specifically rocket-based combined- cycle (RBCC), propulsion systems, typically employ an internal engine flow-path installed primary rocket subsystem. To achieve acceptably short mixing lengths in effecting the "air augmentation" process, a large rocket-exhaust/air interfacial mixing surface is needed. This leads, in some engine design concepts, to a "cluster" of small rocket units, suitably arrayed in the flowpath. To support an early (1964) subscale ground-test of a specific RBCC concept, such a 12-rocket cluster was developed by NASA's Marshall Space Flight Center (MSFC). The small primary rockets used in the cluster assembly were modified versions of an existing small kerosene/oxygen water-cooled rocket engine unit routinely tested at MSFC. Following individual thrust-chamber tests and overall subsystem qualification testing, the cluster assembly was installed at the U. S. Air Force's Arnold Engineering Development Center (AEDC) for RBCC systems testing. (The results of the special air-augmented rocket testing are not covered here.) While this project was eventually successfully completed, a number of hardware integration problems were met, leading to catastrophic thrust chamber failures. The principal "lessons learned" in conducting this early primary rocket subsystem experimental effort are documented here as a basic knowledge-base contribution for the benefit of today's RBCC research and development community.

Impulse control and restrained eating among young women: Evidence for compensatory cortical activation during a chocolate-specific delayed discounting task.

PubMed

Dong, Debo; Wang, Yulin; Jackson, Todd; Chen, Shuaiyu; Wang, Yu; Zhou, Feng; Chen, Hong

2016-10-01

Theory and associated research indicate that people with elevated restrained eating (RE) scores have higher risk for binge eating, future bulimic symptom onset and weight gain. Previous imaging studies have suggested hyper-responsive reward brain area activation in response to food cues contributes to this risk but little is known about associated neural impulse control mechanisms, especially when considering links between depleted cognitive resources related to unsuccessful RE. Towards illuminating this issue, we used a chocolate-specific delayed discounting (DD) task to investigate relations between RE scores, behavior impulsivity, and corresponding neural impulse control correlates in a functional magnetic resonance imaging (fMRI) study of 27 young women. Specifically, participants were required to choose between more immediate, smaller versus delayed, larger hypothetical chocolate rewards following initial consumption of a chocolate. As predicted, RE scores were correlated positively with behavior impulse control levels. More critically, higher RE scores were associated with stronger activation in impulse control region, the dorsal-lateral prefrontal cortex (DLPFC) during the completion of difficult decision trials reflecting higher cognitive demands and resource depletion relative to easy decision trials. Exploratory analyses revealed a positive correlation between RE scores and activity in a reward system hub, the right striatum. Moreover, a positive correlation between left DLPFC and striatum activation was posited to reflect, in part, impulse control region compensation in response to stronger reward signal among women with RE elevations. Findings suggested impulse control lapses may contribute to difficulties in maintaining RE, particularly when cognitive demands are high. Copyright © 2016. Published by Elsevier Ltd.

Domain-Specific Impulsivity in School-Age Children

ERIC Educational Resources Information Center

Tsukayama, Eli; Duckworth, Angela Lee; Kim, Betty

2013-01-01

Impulsivity is a salient individual difference in children with well-established predictive validity for life outcomes. The current investigation proposes that impulsive behaviors vary systematically by domain. In a series of studies with ethnically and socioeconomically diverse samples of middle school students, we find that schoolwork-related…

Improving Inhibitory Control Abilities (ImpulsE)-A Promising Approach to Treat Impulsive Eating?

PubMed

Preuss, Hanna; Pinnow, Marlies; Schnicker, Katja; Legenbauer, Tanja

2017-11-01

Although there is preliminary evidence that inhibitory control training improves impulsive eating, less is known about the effects on eating behaviour and weight loss in clinical samples. Sixty-nine treatment-seeking adults with obesity (binge-eating disorder 33.3%; other specific feeding and eating disorders 40.6%) were randomly blockwise allocated to ImpulsE, an intervention to improve inhibitory control and emotion regulation abilities or a guideline-appropriate cognitive behavioural therapy (CBT)-based treatment as usual. Self-reported and performance-based impulsivity, eating disorder pathology and BMI were compared at baseline (T1), post-treatment (T2) and 1- or 3-month follow-up. ImpulsE led to better food-specific inhibition performance (p = .004), but groups did not differ regarding improvements in global Eating Disorder Examination Questionnaire (EDE-Q) score at T2. At 3-month follow-up, binge eaters benefited most from ImpulsE (p = .028) and completers of ImpulsE demonstrated a significantly greater weight reduction (p = .030). The current findings propose ImpulsE as a promising approach to treat obesity, illustrating acceptability and additional benefits for course of weight. Copyright © 2017 John Wiley & Sons, Ltd and Eating Disorders Association. Copyright © 2017 John Wiley & Sons, Ltd and Eating Disorders Association.

Motor impulsivity during childhood and adolescence: a longitudinal biometric analysis of the go/no-go task in 9- to 18-year-old twins.

PubMed

Bezdjian, Serena; Tuvblad, Catherine; Wang, Pan; Raine, Adrian; Baker, Laura A

2014-11-01

In the present study, we investigated genetic and environmental effects on motor impulsivity from childhood to late adolescence using a longitudinal sample of twins from ages 9 to 18 years. Motor impulsivity was assessed using errors of commission (no-go errors) in a visual go/no-go task at 4 time points: ages 9-10, 11-13, 14-15, and 16-18 years. Significant genetic and nonshared environmental effects on motor impulsivity were found at each of the 4 waves of assessment with genetic factors explaining 22%-41% of the variance within each of the 4 waves. Phenotypically, children's average performance improved across age (i.e., fewer no-go errors during later assessments). Multivariate biometric analyses revealed that common genetic factors influenced 12%-40% of the variance in motor impulsivity across development, whereas nonshared environmental factors common to all time points contributed to 2%-52% of the variance. Nonshared environmental influences specific to each time point also significantly influenced motor impulsivity. Overall, results demonstrated that although genetic factors were critical to motor impulsivity across development, both common and specific nonshared environmental factors played a strong role in the development of motor impulsivity across age. (PsycINFO Database Record (c) 2014 APA, all rights reserved).

Development of an electrostatic propulsion engine using sub-micron powders as the reaction mass

NASA Technical Reports Server (NTRS)

Herbert, F.; Kendall, K. R.

1991-01-01

Asteroid sample return missions would benefit from development of an improved rocket engine. Chemical rockets achieve their large thrust with high mass consumption rate (dm/dt) but low exhaust velocity; therefore, a large fraction of their total mass is fuel. Present day ion thrusters are characterized by high exhaust velocity, but low dm/dt; thus, they are inherently low thrust devices. However, their high exhausy velocity is poorly matched to typical mission requirements and therefore, wastes energy. A better match would be intermediate between the two forms of propulsion. This could be achieved by electrostatically accelerating solid powder grains, raising the possibility that interplanetary material could be processed to use as reaction mass. An experiment to study the charging properties of sub-micron sized powder grains is described. If a suitable material can be identified, then it could be used as the reaction mass in an electrostatic propulsion engine. The experiment employs a time of flight measurement to determine the exhaust velocity (v) of various negatively charged powder grains that were charged and accelerated in a simple device. The purpose is to determine the charge to mass ratio that can be sustained for various substances. In order to be competitive with present day ion thrusters, a specific impulse (v/g) of 3000 to 5000 seconds is required. Preliminary results are presented. More speculatively, there are some mission profiles that would benefit from collection of reaction mass at the remote asteroid site. Experiments that examine the generation of sub-micron clusters by electrostatic self-disruption of geologically derived material are planned.

Sea-Level Flight Demonstration and Altitude Characterization of a LO2/LCH4 Based Accent Propulsion Lander

NASA Technical Reports Server (NTRS)

Collins, Jacob; Hurlbert, Eric; Romig, Kris; Melcher, John; Hobson, Aaron; Eaton, Phil

2009-01-01

A 1,500 lbf thrust-class liquid oxygen (LO2)/Liquid Methane (LCH4) rocket engine was developed and tested at both sea-level and simulated altitude conditions. The engine was fabricated by Armadillo Aerospace (AA) in collaboration with NASA Johnson Space Center. Sea level testing was conducted at Armadillo Aerospace facilities at Caddo Mills, TX. Sea-level tests were conducted using both a static horizontal test bed and a vertical take-off and landing (VTOL) test bed capable of lift-off and hover-flight in low atmosphere conditions. The vertical test bed configuration is capable of throttling the engine valves to enable liftoff and hover-flight. Simulated altitude vacuum testing was conducted at NASA Johnson Space Center White Sands Test Facility (WSTF), which is capable of providing altitude simulation greater than 120,000 ft equivalent. The engine tests demonstrated ignition using two different methods, a gas-torch and a pyrotechnic igniter. Both gas torch and pyrotechnic ignition were demonstrated at both sea-level and vacuum conditions. The rocket engine was designed to be configured with three different nozzle configurations, including a dual-bell nozzle geometry. Dual-bell nozzle tests were conducted at WSTF and engine performance data was achieved at both ambient pressure and simulated altitude conditions. Dual-bell nozzle performance data was achieved over a range of altitude conditions from 90,000 ft to 50,000 ft altitude. Thrust and propellant mass flow rates were measured in the tests for specific impulse (Isp) and C* calculations.

Food-Related Impulsivity in Obesity and Binge Eating Disorder-A Systematic Update of the Evidence.

PubMed

Giel, Katrin E; Teufel, Martin; Junne, Florian; Zipfel, Stephan; Schag, Kathrin

2017-10-27

The specific eating pattern of Binge Eating Disorder (BED) patients has provoked the assumption that BED might represent a phenotype within the obesity spectrum that is characterized by increased impulsivity. Following the guidelines of the PRISMA statement (preferred reporting items for systematic reviews and meta-analyses), we here provide a systematic update on the evidence on food-related impulsivity in obese individuals, with and without BED, as well as normal-weight individuals. We separately analyzed potential group differences in the impulsivity components of reward sensitivity and rash-spontaneous behavior. Our search resulted in twenty experimental studies with high methodological quality. The synthesis of the latest evidence consolidates conclusions drawn in our initial systematic review that BED represents a distinct phenotype within the obesity spectrum that is characterized by increased impulsivity. Rash-spontaneous behavior in general, and specifically towards food, is increased in BED, while food-specific reward sensitivity is also increased in obese individuals without BED, but potentially to a lesser degree. A major next step for research entails the investigation of sub-domains and temporal components of inhibitory control in BED and obesity. Based on the evidence of impaired inhibitory control in BED, affected patients might profit from interventions that address impulsive behavior.

Domain-Specific Relationships in Sexual Measures of Impulsive Behavior.

PubMed

Mahoney, Colin T; Lawyer, Steven R

2018-04-25

Impulsivity is an important construct for understanding sexual behaviors, but behavioral and self-report measures of impulsivity often are not correlated. One possible explanation for this is that there is little shared variance in the measures because behavioral measures index impulsivity by asking questions about monetary preferences, while self-report measures index impulsivity by asking about a broad range of real-world outcomes (including those of a sexual nature) largely unrelated to money-related preferences. Undergraduate students (total N = 105; female n = 77, male n = 28) completed laboratory measures-delay discounting (DD) and probability discounting (PD)-for two different outcomes-money and sexual activity. Participants also completed the Delaying Gratification Inventory (DGI), which measures difficulty with delaying gratification (i.e., impulsivity) across different domains, including money and physical pleasures. Findings indicated that DD and PD for money were not related to any of the DGI subscales. However, DD for sexual activity was significantly related to the DGI Physical Pleasures subscale, but not other subscales. These findings suggest that the relationship between behavioral and self-report measures of impulsive choice may be stronger when both are measuring domain-specific rather than domain-general behavioral patterns, but further research is warranted.

Experimental and simulation study of a Gaseous oxygen/Gaseous hydrogen vortex cooling thrust chamber

NASA Astrophysics Data System (ADS)

Yu, Nanjia; Zhao, Bo; Li, Gongnan; Wang, Jue

2016-01-01

In this paper, RNG k-ε turbulence model and PDF non-premixed combustion model are used to simulate the influence of the diameter of the ring of hydrogen injectors and oxidizer-to-fuel ratio on the specific impulse of the vortex cooling thrust chamber. The simulation results and the experimental tests of a 2000 N Gaseous oxygen/Gaseous hydrogen vortex cooling thrust chamber reveal that the efficiency of the specific impulse improves significantly with increasing of the diameter of the ring of hydrogen injectors. Moreover, the optimum efficiency of the specific impulse is obtained when the oxidizer-to-fuel ratio is near the stoichiometric ratio.

Potential low cost, safe, high efficiency propellant for future space program

NASA Astrophysics Data System (ADS)

Zhou, D.

2005-03-01

Mixtures of nanometer or micrometer sized carbon powder suspended in hydrogen and methane/hydrogen mixtures are proposed as candidates for low cost, high efficiency propellants for future space programs. While liquid hydrogen has low weight and high heat of combustion per unit mass, because of the low mass density the heat of combustion per unit volume is low, and the liquid hydrogen storage container must be large. The proposed propellants can produce higher gross heat combustion with small volume with trade off of some weight increase. Liquid hydrogen can serve as the fluid component of the propellant in the mixtures and thus used by current rocket engine designs. For example, for the same volume a mixture of 5% methane and 95% hydrogen, can lead to an increase in the gross heat of combustion by about 10% and an increase in the Isp (specific impulse) by 21% compared to a pure liquid hydrogen propellant. At liquid hydrogen temperatures of 20.3 K, methane will be in solid state, and must be formed as fine granules (or slush) to satisfy the requirement of liquid propellant engines.

Use of Atomic Fuels for Rocket-Powered Launch Vehicles Analyzed

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan A.

1999-01-01

At the NASA Lewis Research Center, the launch vehicle gross lift-off weight (GLOW) was analyzed for solid particle feed systems that use high-energy density atomic propellants (ref. 1). The analyses covered several propellant combinations, including atoms of aluminum, boron, carbon, and hydrogen stored in a solid cryogenic particle, with a cryogenic liquid as the carrier fluid. Several different weight percents for the liquid carrier were investigated, and the GLOW values of vehicles using the solid particle feed systems were compared with that of a conventional oxygen/hydrogen (O2/H2) propellant vehicle. Atomic propellants, such as boron, carbon, and hydrogen, have an enormous potential for high specific impulse Isp operation, and their pursuit has been a topic of great interest for decades. Recent and continuing advances in the understanding of matter, the development of new technologies for simulating matter at its most basic level, and manipulations of matter through microtechnology and nanotechnology will no doubt create a bright future for atomic propellants and an exciting one for the researchers exploring this technology.

Lunar He-3, fusion propulsion, and space development

NASA Technical Reports Server (NTRS)

Santarius, John F.

1992-01-01

The recent identification of a substantial lunar resource of the fusion energy fuel He-3 may provide the first terrestrial market for a lunar commodity and, therefore, a major impetus to lunar development. The impact of this resource-when burned in D-He-3 fusion reactors for space power and propulsion-may be even more significant as an enabling technology for safe, efficient exploration and development of space. One possible reactor configuration among several options, the tandem mirror, illustrates the potential advantages of fusion propulsion. The most important advantage is the ability to provide either fast, piloted vessels or high-payload-fraction cargo vessels due to a range of specific impulses from 50 sec to 1,000,000 sec at thrust-to-weight ratios from 0.1 to 5x10(exp -5). Fusion power research has made steady, impressive progress. It is plausible, and even probable, that fusion rockets similar to the designs presented here will be available in the early part of the twenty-first century, enabling a major expansion of human presence into the solar system.

Helicon Plasma Injector and Ion Cyclotron Acceleration Development in the VASIMR Experiment

NASA Technical Reports Server (NTRS)

Squire, Jared P.; Chang, Franklin R.; Jacobson, Verlin T.; McCaskill, Greg E.; Bengtson, Roger D.; Goulding, Richard H.

2000-01-01

In the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) radio frequency (rf) waves both produce the plasma and then accelerate the ions. The plasma production is done by action of helicon waves. These waves are circular polarized waves in the direction of the electron gyromotion. The ion acceleration is performed by ion cyclotron resonant frequency (ICRF) acceleration. The Advanced Space Propulsion Laboratory (ASPL) is actively developing efficient helicon plasma production and ICRF acceleration. The VASIMR experimental device at the ASPL is called VX-10. It is configured to demonstrate the plasma production and acceleration at the 10kW level to support a space flight demonstration design. The VX-10 consists of three electromagnets integrated into a vacuum chamber that produce magnetic fields up to 0.5 Tesla. Magnetic field shaping is achieved by independent magnet current control and placement of the magnets. We have generated both helium and hydrogen high density (>10(exp 18) cu m) discharges with the helicon source. ICRF experiments are underway. This paper describes the VX-10 device, presents recent results and discusses future plans.

Propulsion Systems Integration for a `Tractor Beam' Mercury Lightcraft: Liftoff Engine

NASA Astrophysics Data System (ADS)

Myrabo, L. N.

2003-05-01

Described herein is the concept and propulsion systems integration for a revolutionary beam-propelled shuttle called the ``Mercury'' lightcraft - emphasizing the liftoff engine mode. This one-person, ultra-energetic vehicle is designed to ride `tractor beams' into space, transmitted from a future network of satellite solar power stations. The objective is to create a safe, very low cost (e.g., 1000X below chemical rockets) space transportation system for human life, one that is completely `green' and independent of Earth's limited fossil fuel reserves. The lightcraft's airbreathing combined-cycle engine operates in a rotary pulsed detonation mode PDE for lift-offs and landings; at hypersonic speeds it transitions into a magnetohydrodynamic (MHD) slipstream accelerator mode. For the latter, the transatmospheric flight path is momentarily transformed into an extremely long, electromagnetic ``mass-driver'' channel with an effective `fuel' specific impulse in the range of 6000 to 16,000 seconds. These future single-stage-to-orbit, highly-reusuable vehicles will ride ``Highways of Light,'' accelerating at 3 Gs into space, with their throttles just barely beyond `idle' power.

Gas-core reactor power transient analysis

NASA Technical Reports Server (NTRS)

Kascak, A. F.

1972-01-01

The gas core reactor is a proposed device which features high temperatures. It has applications in high specific impulse space missions, and possibly in low thermal pollution MHD power plants. The nuclear fuel is a ball of uranium plasma radiating thermal photons as opposed to gamma rays. This thermal energy is picked up before it reaches the solid cavity liner by an inflowing seeded propellant stream and convected out through a rocket nozzle. A wall-burnout condition will exist if there is not enough flow of propellant to convect the energy back into the cavity. A reactor must therefore operate with a certain amount of excess propellant flow. Due to the thermal inertia of the flowing propellant, the reactor can undergo power transients in excess of the steady-state wall burnout power for short periods of time. The objective of this study was to determine how long the wall burnout power could be exceeded without burning out the cavity liner. The model used in the heat-transfer calculation was one-dimensional, and thermal radiation was assumed to be a diffusion process.

Recent Advancements in Propellant Densification

NASA Technical Reports Server (NTRS)

McNelis, Nancy B.; Tomsik, Thomas M.

1998-01-01

Next-generation launch vehicles demand several technological improvements to achieve lower cost and more reliable access to space. One technology area whose performance gains may far exceed others is densified propellants. The ideal rocket engine propellant is characterized by high specific impulse, high density, and low vapor pressure. A propellant combination of liquid hydrogen and liquid oxygen (LH2/LOX) is one of the highest performance propellants, but LH2 stored at standard conditions has a relatively low density and high vapor pressure. Propellant densification can significantly improve this propellant's properties relative to vehicle design and engine performance. Vehicle performance calculations based on an average of existing launch vehicles indicate that densified propellants may allow an increase in payload mass of up to 5 percent. Since the NASA Lewis Research Center became involved with the National Aerospace Plane program in the 1980's, it has been leading the way in making densified propellants a viable fuel for next-generation launch vehicles. Lewis researchers have been working to provide a method and critical data for continuous production of densified hydrogen and oxygen.

Combustion/particle sizing experiments at the Naval Postgraduate School Combustion Research Laboratory

NASA Technical Reports Server (NTRS)

Powers, John; Netzer, David

1987-01-01

Particle behavior in combustion processes is an active research area at NPS. Currently, four research efforts are being conducted: (1) There is a long standing need to better understand the soot production and combustion processes in gas turbine combustors, both from a concern for improved engine life and to minimize exhaust particulates. Soot emissions are strongly effected by fuel composition and additives; (2) A more recent need for particle sizing/behavior measurements is in the combustor of a solid fuel ramjet which uses a metallized fuel. High speed motion pictures are being used to study rather large burning particles; (3) In solid propellant rocket motors, metals are used to improve specific impulse and/or to provide damping for combustion pressure oscillations. Particle sizing experiments are being conducted using diode arrays to measure the light intensity as a function of scattering angle; (4) Once a good quality hologram is attained, a need exists for obtaining the particle distributions from hologram in a short period of time. A Quantimet 720 Image Analyzer is being used to reconstruct images.

Pulsed Electric Propulsion Thrust Stand Calibration Method

NASA Technical Reports Server (NTRS)

Wong, Andrea R.; Polzin, Kurt A.; Pearson, J. Boise

2011-01-01

The evaluation of the performance of any propulsion device requires the accurate measurement of thrust. While chemical rocket thrust is typically measured using a load cell, the low thrust levels associated with electric propulsion (EP) systems necessitate the use of much more sensitive measurement techniques. The design and development of electric propulsion thrust stands that employ a conventional hanging pendulum arm connected to a balance mechanism consisting of a secondary arm and variable linkage have been reported in recent publications by Polzin et al. These works focused on performing steady-state thrust measurements and employed a static analysis of the thrust stand response. In the present work, we present a calibration method and data that will permit pulsed thrust measurements using the Variable Amplitude Hanging Pendulum with Extended Range (VAHPER) thrust stand. Pulsed thrust measurements are challenging in general because the pulsed thrust (impulse bit) occurs over a short timescale (typically 1 micros to 1 millisecond) and cannot be resolved directly. Consequently, the imparted impulse bit must be inferred through observation of the change in thrust stand motion effected by the pulse. Pulsed thrust measurements have typically only consisted of single-shot operation. In the present work, we discuss repetition-rate pulsed thruster operation and describe a method to perform these measurements. The thrust stand response can be modeled as a spring-mass-damper system with a repetitive delta forcing function to represent the impulsive action of the thruster.

An Analytic Approximation to Very High Specific Impulse and Specific Power Interplanetary Space Mission Analysis

NASA Technical Reports Server (NTRS)

Williams, Craig Hamilton

1995-01-01

A simple, analytic approximation is derived to calculate trip time and performance for propulsion systems of very high specific impulse (50,000 to 200,000 seconds) and very high specific power (10 to 1000 kW/kg) for human interplanetary space missions. The approach assumed field-free space, constant thrust/constant specific power, and near straight line (radial) trajectories between the planets. Closed form, one dimensional equations of motion for two-burn rendezvous and four-burn round trip missions are derived as a function of specific impulse, specific power, and propellant mass ratio. The equations are coupled to an optimizing parameter that maximizes performance and minimizes trip time. Data generated for hypothetical one-way and round trip human missions to Jupiter were found to be within 1% and 6% accuracy of integrated solutions respectively, verifying that for these systems, credible analysis does not require computationally intensive numerical techniques.

Differential contributions of cortical thickness and surface area to trait impulsivity in healthy young adults.

PubMed

Kubera, Katharina M; Schmitgen, Mike M; Maier-Hein, Klaus H; Thomann, Philipp A; Hirjak, Dusan; Wolf, Robert C

2018-05-08

Impulsivity is an essential human personality trait and highly relevant for the development of several mental disorders. There is evidence that impulsivity is heritable, yet little is known about neural correlates reflecting early brain development. Here, we address the question whether motor, attentional and non-planning components, as reflected by the Barratt Impulsiveness Scale (BIS-11), are distinctly associated with cortical thickness and surface area variations in young healthy individuals. We investigated cortical thickness and surface area in 54 healthy volunteers (m/f = 30%/70%; age mean/SD = 24.9/4.02) using structural magnetic resonance imaging at 3 T together with surface-based analysis techniques. Impulsivity was examined on the Barratt impulsiveness scale (BIS-11) and related to the two distinct cortical measurements. Higher BIS-11 total scores were negatively associated with cortical thickness variations in the left lingual gyrus, left superior temporal gyrus, right cuneus, and right superior parietal gyrus (p<0.05 cluster-wise probability [CWP] corrected). Higher BIS-11 nonplanning scores were negatively associated with cortical thickness variations in bilateral pericalcarine gyrus (p<0.05 CWP corr.). In the orbitofrontal cortex motor impulsivity associated cortical thickness differs significantly between male and female. These data suggest distinct neurodevelopmental trajectories underlying impulsivity in healthy subjects. Impulsivity total scores appear to be specifically related to cortical thickness variations, in contrast to variations of cortical surface area. Furthermore, our findings underscores the importance of better characterizing gender-specific structural correlates of impulsivity. Copyright © 2018. Published by Elsevier B.V.

Dysfunctional decision-making in pathological gambling: pattern specificity and the role of impulsivity.

PubMed

Kräplin, Anja; Dshemuchadse, Maja; Behrendt, Silke; Scherbaum, Stefan; Goschke, Thomas; Bühringer, Gerhard

2014-03-30

Dysfunctional decision-making in individuals with pathological gambling (PGs) may result from dominating reward-driven processes, indicated by higher impulsivity. In the current study we examined (1) if PGs show specific decision-making impairments related to dominating reward-driven processes rather than to strategic planning deficits and (2) whether these impairments are related to impulsivity. Nineteen PGs according to DSM-IV and 19 matched control subjects undertook the Cambridge Gambling Task (CGT) to assess decision-making. The delay discounting paradigm (DDP) as well as the UPPS Impulsive Behavior Scale (measuring urgency, premeditation, perseverance and sensation seeking) were administered as multidimensional measures of impulsivity. Results revealed that (1) PGs exhibited higher risk seeking and an immediate reward focus in the CGT and, in contrast, comparable strategic planning to the control group. (2) Decision-making impairments were related to more severe delay discounting and, specifically, to increased urgency and less premeditation. Our findings suggest (1) the necessity to disentangle decision-making components in order to improve etiological models of PGs, and (2) that urgency and premeditation are specifically related to disadvantageous decision-making and should be tackled in intervention strategies focusing on emotion tolerance and control strategies. Copyright © 2014 Elsevier Ireland Ltd. All rights reserved.

Impulsivity and Gambling Type Among Treatment-Seeking Disordered Gamblers: An Explorative Study.

PubMed

Lutri, Vittorio; Soldini, Emiliano; Ronzitti, Silvia; Smith, Neil; Clerici, Massimo; Blaszczynski, Alex; Bowden-Jones, Henrietta

2018-03-03

Several studies have found that certain traits of impulsivity are associated with gambling disorder, and influence its severity. Furthermore, it has been suggested that some forms of gambling, particularly electronic gambling machines, are particularly widespread among pathological gamblers. In the present, exploratory study, we aim to clarify the role played by impulsivity in influencing the choice of specific gambling activities, by examining the relation between individual dimensions of impulsivity, and the choice of specific gambling activities in a clinical population. 100 consecutively admitted pathological gamblers at the National Problem Gambling Clinic in London (UK) in 2014 were administered the UPPS-P and BIS-11 impulsivity questionnaires, the Problem Gambling Severity Index, and underwent a structured interview concerning their gambling activities in the month and year prior to assessment. The correlation between individual gambling activities and impulsivity dimensions was analyzed both at a bivariate level, and using logistic regression. We found a significant correlation between Negative Urgency, Motor impulsivity and low-stakes machine gambling on multivariate analysis. Negative urgency (i.e. the tendency to act impulsively in response to negative affect), and Motor impulsivity (a tendency to rash action and restlessness) might be mediating factors in the choice of electronic gambling machines, particularly among patients whose gambling is escape-oriented. Structural and situational characteristics of gambling machines, particularly the widespread availability of low-stakes-rather than high-stakes-gaming machines, might concur to the choice of this form of gambling among individuals who present higher negative urgency and restlessness.

Validity of proposed DSM-5 ADHD impulsivity symptoms in children.

PubMed

Ünsel Bolat, Gül; Ercan, Eyüp Sabri; Salum, Giovanni Abrahão; Bilaç, Öznur; Massuti, Rafael; Uysal Özaslan, Taciser; Bolat, Hilmi; Rohde, Luis Augusto

2016-10-01

The American Psychiatric Association (APA) working group on Attention-Deficit/Hyperactivity Disorder (ADHD) proposed the inclusion of four new impulsivity symptoms. However, they were not included in DSM-5 due to the lack of sufficient evidence. The aim of this study is to investigate the performance of the proposed four ADHD impulsivity symptoms with respect to: (a) ADHD factor structure; (b) performance in predicting clinical impairment; (c) specificity for ADHD diagnosis and (d) best symptomatic threshold to predict clinical impairment. The sample comprised 416 children (31 ADHD subjects according to both DSM-IV and proposed DSM-5, 20 ADHD subjects according to just one diagnostic system and 365 controls) from 12 schools. Diagnoses were derived using semi-structured interviews and ADHD rating scales. Results from confirmatory factor analysis indicate that addition of the four new impulsivity items provided a slightly better factor structure if compared to models including only 18 items. Regression analyses showed that only one of the new impulsivity symptoms (impatient) was part of the list of best predictors of impairment. None of the four new impulsivity items was specifically associated with ADHD diagnosis. The best cutoff point in the hyperactivity/impulsivity dimension for predicting impairment did not change significantly. Overall, our findings suggest that the determination on how to best capture impulsivity dimension as part of the ADHD construct needs more investigation and that there is not enough evidence to include these four assessed impulsivity symptoms as part of the ADHD criteria.

Options for human ``return to the moon'' using tomorrow's SSTO, ISRU, and LOX-augmented NTR technologies

NASA Astrophysics Data System (ADS)

Borowski, Stanley K.

1996-03-01

The feasibility of conducting human missions to the Moon is examined assuming the use of three ``high leverage'' technologies: (1) a single-stage-to-orbit (SSTO) launch vehicle, (2) ``in-situ'' resource utilization (ISRU)—specifically ``lunar-derived'' liquid oxygen (LUNOX), and (3) LOX-augmented nuclear thermal rocket (LANTR) propulsion. Lunar transportation system elements consisting of a LANTR-powered lunar transfer vehicle (LTV) and a chemical propulsion lunar landing/Earth return vehicle (LERV) are configured to fit within the ``compact'' dimensions of the SSTO cargo bay (diameter: 4.6 m/length: 9.0 m) while satisfying an initial mass in low Earth orbit (IMLEO) limit of ˜60 t (3 SSTO launches). Using ˜8 t of LUNOX to ``reoxidize'' the LERV for a ``direct return'' flight to Earth reduces its size and mass allowing delivery to LEO on a single 20 t SSTO launch. Similarly, the LANTR engine's ability to operate at any oxygen/hydrogen mixture ratio from 0 to 7 with high specific impulse (˜940 to 515 s) is exploited to reduce hydrogen tank volume, thereby improving packaging of the LANTR LTV's ``propulsion'' and ``propellant modules''. Expendable and reusable, piloted and cargo missions and vehicle designs are presented along with estimates of LUNOX production required to support the different mission modes.

Dynamic response of laser ablative shock waves from coated and uncoated amorphous Boron nanoparticles

NASA Astrophysics Data System (ADS)

Paturi, Prem Kiran; Chelikani, Leela; Pinnoju, Venkateshwarlu; Verma, Pankaj; Singh, Raja V.; Acrhem Collaboration; Hemrl Collaboration

2015-06-01

Nanoparticles (NP) improve the performance of solid rocket motors with increased burning rate and lower ignition threshold owing to their larger surface area. We present spatio-temporal evolution of laser ablative shock waves (LASWs) from compacted amorphous Boron (B) and Lithium Fluoride coated Boron (LiF-B) of 70-110nm sizes that were compacted to form pellets. Thickness of the LiF coating is 5.5 +/- 1 nm in LiF-B. Laser pulses from second harmonic of Nd:YAG laser (532 nm, 7 ns) are used to generate LASWs expanding in ambient air. The precise time of energy release from the pellets under extreme ablative pressures is studied using shadowgraphy with a temporal resolution of 1.5 ns. Different nature of the shock front (SF) following Sedov-Taylor theory, before and after detachment, indicated two specific time dependent stages of energy release. From the position of SF, velocity behind the SF, similar to that of exhaust velocity is measured. Specific impulse of 241 +/- 5 and 201 +/- 4 sec for LiF-B and B, respectively, at a delay of 0.8 μs from shock inducing laser pulse makes them potential candidates for laser based micro thruster applications. The work is supported by Defence Research and Developement Organization, India through Grants-in-Aid Program.

The advisability of prototypic testing for space nuclear systems

NASA Astrophysics Data System (ADS)

Lenard, Roger X.

2005-07-01

From October 1987 until 1993, the US Department of Defense conducted the Space Nuclear Thermal Propulsion program. This program's objective was to design and develop a high specific impulse, high thrust-to-weight nuclear thermal rocket engine for upper stage applications. The author was the program manager for this program until 1992. Numerous analytical, programmatic and experimental results were generated during this period of time. This paper reviews the accomplishments of the program and highlights the importance of prototypic testing for all aspects of a space nuclear program so that a reliable and safe system compliant with all regulatory requirements can be effectively engineered. Specifically, the paper will recount how many non-prototypic tests we performed only to have more representative tests consistently generate different results. This was particularly true in area of direct nuclear heat generation. As nuclear tests are generally much more expensive than non-nuclear tests, programs attempt to avoid such tests in favor of less expensive non-nuclear tests. Each time this approach was followed, the SNTP program found these tests to not be verified by nuclear heated testing. Hence the author recommends that wherever possible, a spiral development approach that includes exploratory and confirmatory experimental testing be employed to ensure a viable design.

Advanced concepts. [specific impulse, mass drivers, electromagnetic launchers, and the rail gun

NASA Technical Reports Server (NTRS)

Banks, B. A.

1980-01-01

The relative strengths of those interactions which enable propulsive forces are listed as well as the specific impulse of various propellants. Graphics show the linear synchronous motor of the mass driver, the principle of the direct current electromagnetic launcher, and the characteristics of the rail gun.

The evolutionary development of high specific impulse electric thruster technology

NASA Technical Reports Server (NTRS)

Sovey, James S.; Hamley, John A.; Patterson, Michael J.; Rawlin, Vincent K.; Myers, Roger M.

1992-01-01

Electric propulsion flight and technology demonstrations conducted primarily by Europe, Japan, China, the U.S., and the USSR are reviewed. Evolutionary mission applications for high specific impulse electric thruster systems are discussed, and the status of arcjet, ion, and magnetoplasmadynamic thrusters and associated power processor technologies are summarized.

Impulsivity in Alcohol-Dependent Patients with and without ADHD: The Role of Atomoxetine.

PubMed

Coppola, Maurizio; Mondola, Raffaella

2018-06-07

Attention deficit hyperactivity disorder (ADHD) is a neurodevelopmental disorder characterized by symptoms of hyperactivity, inattention, and impulsivity. As shown in the literature, this disorder is frequently associated with alcohol and drug abuse. Patients affected by ADHD show high levels of impulsivity and sensation seeking. These characteristics can significantly increase the risk of alcohol abuse, which is itself a clinical condition associated with high levels of impulsivity. Clinical studies suggest that atomoxetine is effective and safe in patients affected by both ADHD and alcohol dependence; however, information focused specifically on impulsivity is very limited. In an open-label study, we evaluated the difference in impulsivity level between alcohol-dependent patients with and without a diagnosis of ADHD. Furthermore, we hypothesized that, in patients with ADHD, atomoxetine could reduce the impulsivity trait.

Study of solid rocket motors for a space shuttle booster. Appendix B: Prime item development specification

NASA Technical Reports Server (NTRS)

1972-01-01

The specifications for the performance, design, development, and test requirements of the P2-156, S3-156, and S6-120 space shuttle booster solid rocket motors are presented. The applicable documents which form a part of the specifications are listed.

Genetics of impulsive behaviour

PubMed Central

Bevilacqua, Laura; Goldman, David

2013-01-01

Impulsivity, defined as the tendency to act without foresight, comprises a multitude of constructs and is associated with a variety of psychiatric disorders. Dissecting different aspects of impulsive behaviour and relating these to specific neurobiological circuits would improve our understanding of the etiology of complex behaviours for which impulsivity is key, and advance genetic studies in this behavioural domain. In this review, we will discuss the heritability of some impulsivity constructs and their possible use as endophenotypes (heritable, disease-associated intermediate phenotypes). Several functional genetic variants associated with impulsive behaviour have been identified by the candidate gene approach and re-sequencing, and whole genome strategies can be implemented for discovery of novel rare and common alleles influencing impulsivity. Via deep sequencing an uncommon HTR2B stop codon, common in one population, was discovered, with implications for understanding impulsive behaviour in both humans and rodents and for future gene discovery. PMID:23440466

MHD Simulation of Magnetic Nozzle Plasma with the NIMROD Code: Applications to the VASIMR Advanced Space Propulsion Concept

NASA Astrophysics Data System (ADS)

Tarditi, Alfonso G.; Shebalin, John V.

2002-11-01

A simulation study with the NIMROD code [1] is being carried on to investigate the efficiency of the thrust generation process and the properties of the plasma detachment in a magnetic nozzle. In the simulation, hot plasma is injected in the magnetic nozzle, modeled as a 2D, axi-symmetric domain. NIMROD has two-fluid, 3D capabilities but the present runs are being conducted within the MHD, 2D approximation. As the plasma travels through the magnetic field, part of its thermal energy is converted into longitudinal kinetic energy, along the axis of the nozzle. The plasma eventually detaches from the magnetic field at a certain distance from the nozzle throat where the kinetic energy becomes larger than the magnetic energy. Preliminary NIMROD 2D runs have been benchmarked with a particle trajectory code showing satisfactory results [2]. Further testing is here reported with the emphasis on the analysis of the diffusion rate across the field lines and of the overall nozzle efficiency. These simulation runs are specifically designed for obtaining comparisons with laboratory measurements of the VASIMR experiment, by looking at the evolution of the radial plasma density and temperature profiles in the nozzle. VASIMR (Variable Specific Impulse Magnetoplasma Rocket, [3]) is an advanced space propulsion concept currently under experimental development at the Advanced Space Propulsion Laboratory, NASA Johnson Space Center. A plasma (typically ionized Hydrogen or Helium) is generated by a RF (Helicon) discharge and heated by an Ion Cyclotron Resonance Heating antenna. The heated plasma is then guided into a magnetic nozzle to convert the thermal plasma energy into effective thrust. The VASIMR system has no electrodes and a solenoidal magnetic field produced by an asymmetric mirror configuration ensures magnetic insulation of the plasma from the material surfaces. By powering the plasma source and the heating antenna at different levels it is possible to vary smoothly of the thrust-to-specific impulse ratio while maintaining maximum power utilization. [1] http://www.nimrodteam.org [2] A. V. Ilin et al., Proc. 40th AIAA Aerospace Sciences Meeting, Reno, NV, Jan. 2002 [3] F. R. Chang-Diaz, Scientific American, p. 90, Nov. 2000

Solid Rocket Launch Vehicle Explosion Environments

NASA Technical Reports Server (NTRS)

Richardson, E. H.; Blackwood, J. M.; Hays, M. J.; Skinner, T.

2014-01-01

Empirical explosion data from full scale solid rocket launch vehicle accidents and tests were collected from all available literature from the 1950s to the present. In general data included peak blast overpressure, blast impulse, fragment size, fragment speed, and fragment dispersion. Most propellants were 1.1 explosives but a few were 1.3. Oftentimes the data from a single accident was disjointed and/or missing key aspects. Despite this fact, once the data as a whole was digitized, categorized, and plotted clear trends appeared. Particular emphasis was placed on tests or accidents that would be applicable to scenarios from which a crew might need to escape. Therefore, such tests where a large quantity of high explosive was used to initiate the solid rocket explosion were differentiated. Also, high speed ground impacts or tests used to simulate such were also culled. It was found that the explosions from all accidents and applicable tests could be described using only the pressurized gas energy stored in the chamber at the time of failure. Additionally, fragmentation trends were produced. Only one accident mentioned the elusive "small" propellant fragments, but upon further analysis it was found that these were most likely produced as secondary fragments when larger primary fragments impacted the ground. Finally, a brief discussion of how this data is used in a new launch vehicle explosion model for improving crew/payload survival is presented.

ELF Sferics Produced by Rocket-Triggered Lightning and Observed at Great Distances

NASA Astrophysics Data System (ADS)

Dupree, N. A.; Moore, R. C.; Fraser-Smith, A. C.

2013-12-01

Experimental observations of ELF radio atmospherics produced by rocket-triggered lightning flashes are used to analyze Earth-ionosphere waveguide excitation and propagation characteristics as a function of return stroke. Rocket-triggered lightning experiments are performed at the International Center for Lightning Research and Testing (ICLRT) located at Camp Blanding, Florida. Long-distance ELF observations are performed in California, Greenland, and Antarctica, although this work focuses on observations performed in Greenland. The lightning current waveforms directly measured at the base of the lightning channel (at the ICLRT) are used together with the Long Wavelength Propagation Capability (LWPC) code to predict the sferic waveform observed at the receiver locations under various ionospheric conditions. LWPC was developed by the Naval Ocean Systems Center over a period of many years. It is an inherently narrowband propagation code that has been modified to predict the broadband response of the Earth-ionosphere waveguide to an impulsive lightning flash while preserving the ability of LWPC to account for an inhomogeneous waveguide. This paper critically compares observations with model predictions, and in particular analyzes Earth-ionosphere waveguide excitation as a function of return stroke. The ability to infer source characteristics using observations at great distances may prove to greatly enhance the understanding of lightning processes that are associated with the production of transient luminous events (TLEs) as well as other ionospheric effects associated with lightning.

Concept Assessment of a Fission Fragment Rocket Engine (FFRE) Propelled Spacecraft

NASA Technical Reports Server (NTRS)

Werka, Robert; Clark, Rod; Sheldon, Rob; Percy, Tom

2012-01-01

The March, 2012 issue of Aerospace America stated that ?the near-to-medium prospects for applying advanced propulsion to create a new era of space exploration are not very good. In the current world, we operate to the Moon by climbing aboard a Carnival Cruise Lines vessel (Saturn 5), sail from the harbor (liftoff) shedding whole decks of the ship (staging) along the way and, having reached the return leg of the journey, sink the ship (burnout) and return home in a lifeboat (Apollo capsule). Clearly this is an illogical way to travel, but forced on Explorers by today's propulsion technology. However, the article neglected to consider the one propulsion technology, using today's physical principles that offer continuous, substantial thrust at a theoretical specific impulse of 1,000,000 sec. This engine unequivocally can create a new era of space exploration that changes the way spacecraft operate. Today's space Explorers could travel in Cruise Liner fashion using the technology not considered by Aerospace America, the novel Dusty Plasma Fission Fragment Rocket Engine (FFRE). This NIAC study addresses the FFRE as well as its impact on Exploration Spacecraft design and operation. It uses common physics of the relativistic speed of fission fragments to produce thrust. It radiatively cools the fissioning dusty core and magnetically controls the fragments direction to practically implement previously patented, but unworkable designs. The spacecraft hosting this engine is no more complex nor more massive than the International Space Station (ISS) and would employ the successful ISS technology for assembly and check-out. The elements can be lifted in "chunks" by a Heavy Lift Launcher. This Exploration Spacecraft would require the resupply of small amounts of nuclear fuel for each journey and would be an in-space asset for decades just as any Cruise Liner on Earth. This study has synthesized versions of the FFRE, integrated one concept onto a host spacecraft designed for manned travel to Jupiter's moon, Callisto, and assessed that round trip journey. This engine, although unoptimized, produced 10 pounds force of thrust at a delivered specific impulse of 527,000 seconds for the entire 15-year mission while providing enormous amounts of electrical power to the spacecraft. A payload of 60 metric tons, included in the 300 metric ton vehicle, was carried to Callisto and back; the propellant tanks holding the 4 metric tons of fuel were not jettisoned in the process. The study concluded that the engine and spacecraft are within today's technology, could be built, tested, launched on several SLS (Space Launch System) (or similar) launchers, integrated, checked out, moved to an in-space base such as at a Lagrange point and operated for decades.

Analysis of a Radioisotope Thermal Rocket Engine

NASA Technical Reports Server (NTRS)

Machado-Rodriguez, Jonathan P.; Landis, Geoffrey A.

2016-01-01

The Triton Hopper is a concept for a global hopper vehicle which uses a radioisotope rocket engine and In-situ propellant acquisition to explore the surface of Neptune's moon, Triton. The current Triton Hopper concept stores heated Nitrogen in a spherical tank to be used as the propellant. The aim of the research was to investigate the benefits of storing propellant at ambient temperature and heating it through the use of a thermal block during engine operation, as opposed to storing gas at a high temperature. Lithium, Lithium Fluoride and Beryllium were considered as possible materials for the thermal block. A heat energy analysis indicated that a lithium thermal mass would provide the highest heat energy for a temperature change from 900 Celsius to -100 Celsius. A heat transfer analysis was performed for Nitrogen at -100 Celsius flowing through 1000 passages inside a 1kg lithium thermal block at a temperature of 900 Celsius. The system was analyzed as turbulent flow through a tube with constant surface temperature. The analysis indicated that the propellant reached a maximum temperature of 877 Celsius before entering the nozzle. At this exit temperature, the average specific impulse [I(sub sp)] of the engine was determined to be 157s. Previous studies for the stored heated gas concept suggest that the engine would have an average I(sub sp) of approximately 52s. Thus, the use of a thermal block concept results in a 200 percent engine performance increase. In addition, a tank sizing study was performed to determine if the concept is feasible in terms of mass requirements. The mass for a spherical carbon fiber COPV storing 35kg of nitrogen at an initial temperature of -100 Celsius and a pressure of 1000psia, was determined to be 7.2kg. The specific impulse analysis indicated that the maximum engine performance is obtained for a mass ratio of 5kg of Nitrogen per every 1kg of lithium thermal mass. Thus for 35kg of Nitrogen the total thermal mass would be 7kg. This brings the total mass of the system to 49.2.kg which is less than the 56kg landing payload capacity of the Triton Hopper. Finally, an insulation analysis using 10mm of MLI insulation indicated that a total of 22 watts of heat are lost to the environment. With the heat loss known, the power required to heat the thermal mass to 900 Celsius in 24 days was determined to be 2.15 watts. The study's results allowed us to conclude that the thermal mass concept is the better option due to the performance increase provided, the low power requirement and its compliance with the landing mass requirement of the Triton Hopper.

Long life monopropellant hydrazine thruster evaluation for Space Station Freedom application

NASA Technical Reports Server (NTRS)

Popp, Christopher G.; Henderson, John B.

1991-01-01

In support of propulsion system thruster development activity for Space Station Freedom (SSF), NASA Johnson Space Center (JSC) is conducting a hydrazine thruster technology demonstration program. The goal of this program is to identify impulse life capability of state-of-the-art long life hydrazine thrusters nominally rated for 50 pounds thrust at 300 psia supply pressure. The SSF propulsion system requirement for impulse life of this thruster class is 1.5 million pound-seconds, corresponding to a throughput of approximately 6400 pounds of propellant, with a high performance (234 pound-seconds per propellant pound). Long life thrusters were procured from Hamilton Standard, The Marquardt Company, and Rocket Research Company. Testing has initiated on the thruster designs to identify life while simulating expected thruster firing duty cycles and durations for SSF using monopropellant grade hydrazine. This paper presents a review of the SSF propulsion system and requirements as applicable to hydrazine thrusters, the three long life thruster designs procured by JSC and the resultant acceptance test data for each thruster, and the JSC test plan and facility.

Mission Analysis for High Specific Impulse Deep Space Exploration

NASA Technical Reports Server (NTRS)

Adams, Robert B.; Polsgrove, Tara; Brady, Hugh J. (Technical Monitor)

2002-01-01

This paper describes trajectory calculations for high specific impulse engines. Specific impulses on the order of 10,000 to 100,000 sec are predicted in a variety of fusion powered propulsion systems. This paper and its companion paper seek to build on analyses in the literature to yield an analytical routine for determining time of flight and payload fraction to a predetermined destination. The companion paper will compare the results of this analysis to the trajectories determined by several trajectory codes. The major parameters that affect time of flight and payload fraction will be identified and their sensitivities quantified. A review of existing fusion propulsion concepts and their capabilities will also be tabulated.

Current evaluation of the tripropellant concept

NASA Technical Reports Server (NTRS)

Zurawski, R. L.

1986-01-01

An analytical study was conducted to determine the specific-impulse advantages of adding metals to conventional liquid-bipropellant systems. These tripropellant systems theoretically offer higher specific impulse and increased propellant density compared with bipropellant systems. Metals considered were Be, Li, and Al. Bipropellant systems were H2/O2, N2H4/N2O4, RP-1/O2, and H2/F2. Thermochemical calculations were performed for sea-level expansion from 6.895-MN/sq. m. (1000-psia) chamber pressure over a wide range of mixture ratios and propellant compositions. Three-dimensional plots characterize the specific impulse of each tripropellant system. Technology issues pertinent to metallized propellant systems are discussed.

Relationship between anhedonia and impulsivity in schizophrenia, major depression and schizoaffective disorder.

PubMed

Amr, Mostafa; Volpe, Fernando Madalena

2013-12-01

Anhedonia and impulsivity are prominent symptoms of many psychiatric disorders and may indicate worse prognosis, notably in schizophrenia and major depression. Despite the convergence of negative outcomes from both dimensions, the relationship between anhedonia and impulsivity in psychiatric disorders has been seldom directly assessed. The objective of the present study is to examine the correlations between anhedonia and impulsivity in three diagnostic groups: major depression, schizophrenia and schizoaffective disorder. 121 outpatients (Mansoura University Hospital, Egypt) with major depressive disorder (N=29), schizophrenia (N=59), and schizoaffective disorder (N=33), were assessed and responded to the Beck Depression Inventory, Barrat's Impulsivity Scale-11, and Chapman's Social and Physical Anhedonia Scales. Physical and social anhedonia scores were negatively correlated to impulsivity scores in major depression patients. Conversely, higher scores in physical and social anhedonia predicted higher impulsivity scores in schizophrenia. No correlations between impulsivity and anhedonia were evidenced among schizoaffectives. The relationship between self-reported physical and social anhedonia and impulsivity is diagnosis-specific. Copyright © 2013 Elsevier B.V. All rights reserved.

Space Travel is Utter Bilge: Early Ideas on Interplanetary Exploration

NASA Astrophysics Data System (ADS)

Yeomans, D. K.

2003-12-01

Until a few decades ago, interplanetary travel was the stuff of dreams but the dreamers often turned out to be farsighted while the predictions of some eminent scientists were far too conservative. The prescient dreamers include the Russian schoolteacher, Konstanin Tsiolkovsky who, in 1883, was the first to note that only rockets could serve the needs of space travel. In 1923, Herman Oberth published a treatise discussing various aspects of interplanetary travel including the impulse necessary to escape the Earth's gravitational pull. In his spare time, a German civil engineer, Walter Hohmann, established in 1925 that the optimal energy transfer orbit between planets is an ellipse that is tangent to the orbits of both bodies. Four year later, an Austrian army officer, Hermann Potocnik outlined the benefits of space stations including those in geosynchronous orbits. Whereas Tsiolkovsky, Oberth, Hohmann, and Potocnik provided ideas and theories, the American, Robert H. Goddard, was testing liquid fueled rockets by as early as 1925. By the time he was finished in 1941, Goddard flew liquid fueled rockets that reached speeds of 700 mph and altitudes above 8,000 feet. In direct contrast to the advances by these mostly amateur engineers, many respected authorities scoffed at space travel because of the insurmountable technological difficulties. One year prior to the launch of Sputnik, the British Astronomer Royal, Sir Richard Wooley, declared, "space travel is utter bilge." While the theories of space travel were well developed by the late 1920's, space travel technology was still a poorly funded, mostly amateur, endeavor until the German army hired Oberth's student, Werner von Braun, and others to develop long range rockets for military purposes. In the early 1940's, Von Braun's team developed the rocket propulsion and guidance systems that would one day form the basis of the American space program.

Characterization of advanced electric propulsion systems

NASA Technical Reports Server (NTRS)

Ray, P. K.

1982-01-01

Characteristics of several advanced electric propulsion systems are evaluated and compared. The propulsion systems studied are mass driver, rail gun, MPD thruster, hydrogen free radical thruster and mercury electron bombardment ion engine. These are characterized by specific impulse, overall efficiency, input power, average thrust, power to average thrust ratio and average thrust to dry weight ratio. Several important physical characteristics such as dry system mass, accelerator length, bore size and current pulse requirement are also evaluated in appropriate cases. Only the ion engine can operate at a specific impulse beyond 2000 sec. Rail gun, MPD thruster and free radical thruster are currently characterized by low efficiencies. Mass drivers have the best performance characteristics in terms of overall efficiency, power to average thrust ratio and average thrust to dry weight ratio. But, they can only operate at low specific impulses due to large power requirements and are extremely long due to limitations of driving current. Mercury ion engines have the next best performance characteristics while operating at higher specific impulses. It is concluded that, overall, ion engines have somewhat better characteristics as compared to the other electric propulsion systems.

46 CFR 160.036-3 - Materials, workmanship, construction and performance requirements.

Code of Federal Regulations, 2010 CFR

2010-10-01

...) EQUIPMENT, CONSTRUCTION, AND MATERIALS: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Hand-Held Rocket... performance requirements. (a) Materials. The materials used in handheld rocket-propelled parachute red flare... protected against corrosion. (b) Workmanship. Handheld rocket-propelled parachute red flare distress signals...

46 CFR 160.036-3 - Materials, workmanship, construction and performance requirements.

Code of Federal Regulations, 2012 CFR

2012-10-01

...) EQUIPMENT, CONSTRUCTION, AND MATERIALS: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Hand-Held Rocket... performance requirements. (a) Materials. The materials used in handheld rocket-propelled parachute red flare... protected against corrosion. (b) Workmanship. Handheld rocket-propelled parachute red flare distress signals...

46 CFR 160.036-3 - Materials, workmanship, construction and performance requirements.

Code of Federal Regulations, 2014 CFR

2014-10-01

...) EQUIPMENT, CONSTRUCTION, AND MATERIALS: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Hand-Held Rocket... performance requirements. (a) Materials. The materials used in handheld rocket-propelled parachute red flare... protected against corrosion. (b) Workmanship. Handheld rocket-propelled parachute red flare distress signals...

46 CFR 160.036-3 - Materials, workmanship, construction and performance requirements.

Code of Federal Regulations, 2013 CFR

2013-10-01

...) EQUIPMENT, CONSTRUCTION, AND MATERIALS: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Hand-Held Rocket... performance requirements. (a) Materials. The materials used in handheld rocket-propelled parachute red flare... protected against corrosion. (b) Workmanship. Handheld rocket-propelled parachute red flare distress signals...

46 CFR 160.036-3 - Materials, workmanship, construction and performance requirements.

Code of Federal Regulations, 2011 CFR

2011-10-01

...) EQUIPMENT, CONSTRUCTION, AND MATERIALS: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Hand-Held Rocket... performance requirements. (a) Materials. The materials used in handheld rocket-propelled parachute red flare... protected against corrosion. (b) Workmanship. Handheld rocket-propelled parachute red flare distress signals...

Impulsivity, aggression and suicidal behavior in unipolar and bipolar disorders.

PubMed

Perroud, Nader; Baud, Patrick; Mouthon, Dominique; Courtet, Philippe; Malafosse, Alain

2011-11-01

Predictors of suicidal behaviors (SB) in bipolar (BD) and major depressive disorder (MDD) patients are poorly understood. It has been recognized that behavioral dysregulation characterizes SB with traits of impulsivity and aggression being particularly salient. However, little is known about how these traits are segregated among mood disorder patients with and without a history of suicide attempt (SA). This article aims to compare impulsivity and aggression between 143 controls, 138 BD and 186 MDD subjects with or without a history of SA. BD and MDD patients showed higher impulsivity scores (BIS-10 = 57.9 vs. 44.7, p < 0.0001) and more severe lifetime aggression than controls (Lifetime History of Aggression = 7.3 vs. 3.9, p < 0.0001). Whereas impulsivity helped to distinguish MDD subjects without a history of SA from those with such a history, this was not the case in BD subjects where no difference in impulsive traits was observed between BD without and with history of SA (57.2 vs. 63.2 for BIS-10; p = 0.259). Impulsive and aggressive traits were strongly correlated in suicide attempters (independently of the diagnosis) but not in non-suicide attempters. Dimensional traits were not characterized at different stages of illness. Impulsivity, as a single trait, may be a reliable suicide risk marker in MDD but not in BD patients, and its strong correlation with aggressive traits seems specifically related to SB. Our study therefore suggests that the specific dimension of impulsive aggression should be systematically assessed in mood disorder patients to address properly their suicidal risk. Copyright © 2011 Elsevier B.V. All rights reserved.

Active space debris removal by using laser propulsion

NASA Astrophysics Data System (ADS)

Rezunkov, Yu. A.

2013-03-01

At present, a few projects on the space debris removal by using highpower lasers are developed. One of the established projects is the ORION proposed by Claude Phipps from Photonics Associates Company and supported by NASA (USA) [1]. But the technical feasibility of the concept is limited by sizes of the debris objects (from 1 to 10 cm) because of a small thrust impulse generated at the laser ablation of the debris materials. At the same time, the removal of rocket upper stages and satellites, which have reached the end of their lives, has been carried out only in a very small number of cases and most of them remain on the Low Earth Orbits (LEO). To reduce the amount of these large-size objects, designing of space systems allowing deorbiting upper rocket stages and removing large-size satellite remnants from economically and scientifically useful orbits to disposal ones is considered. The suggested system is based on high-power laser propulsion. Laser-Orbital Transfer Vehicle (LOTV) with the developed aerospace laser propulsion engine is considered as applied to the problem of mitigation of man-made large-size space debris in LEO.

Impulsivity makes more susceptible to overeating after contextual appetitive conditioning.

PubMed

van den Akker, Karolien; Jansen, Anita; Frentz, Florentine; Havermans, Remco C

2013-11-01

Animals can learn that specific contexts are associated with important biological events such as food intake through classical conditioning. Very few studies suggest this is also possible in humans and contextual appetitive conditioning might even be a main determinant of habitual overeating in vulnerable humans. A Virtual Reality laboratory was used to test whether humans show conditioned responding (increased food desires and expectations, increased salivation and increased food intake) to a specific context after repeated pairings of this context with intake. It was also examined whether the personality trait impulsivity strengthens this contextual appetitive conditioning. Conditioned context-induced reactivity was indeed demonstrated and impulsivity predicted increased intake in only the intake-associated context. It is concluded that humans easily learn desires to eat in intake-related environments. The data also suggest that in particular more impulsive people are vulnerable for conditioned context-induced overeating. This relatively easy learning of associations between specific contexts and intake might stimulate habitual overeating and contribute to increased obesity prevalence. Copyright © 2013 Elsevier Ltd. All rights reserved.

Impulsivity in the supermarket. Responses to calorie taxes and subsidies in healthy weight undergraduates.

PubMed

Giesen, Janneke C A H; Havermans, Remco C; Nederkoorn, Chantal; Jansen, Anita

2012-02-01

The present study investigated the effect of taxing high-energy dense products and subsidizing low-energy dense products on changes in calorie consumption. More specifically, we hypothesized that 'more impulsive' individuals were less influenced by such pricing strategies compared to 'less impulsive' individuals. Contrary to our hypothesis, results showed that 'more impulsive' individuals adjusted their calorie consumption with regard to price changes whereas 'less impulsive' participants were less influenced by price changes. Furthermore, taxing high-energy dense products was more successful in reducing calorie consumption than subsidizing low-energy dense products. Copyright © 2011 Elsevier Ltd. All rights reserved.

Probabilistic failure assessment with application to solid rocket motors

NASA Technical Reports Server (NTRS)

Jan, Darrell L.; Davidson, Barry D.; Moore, Nicholas R.

1990-01-01

A quantitative methodology is being developed for assessment of risk of failure of solid rocket motors. This probabilistic methodology employs best available engineering models and available information in a stochastic framework. The framework accounts for incomplete knowledge of governing parameters, intrinsic variability, and failure model specification error. Earlier case studies have been conducted on several failure modes of the Space Shuttle Main Engine. Work in progress on application of this probabilistic approach to large solid rocket boosters such as the Advanced Solid Rocket Motor for the Space Shuttle is described. Failure due to debonding has been selected as the first case study for large solid rocket motors (SRMs) since it accounts for a significant number of historical SRM failures. Impact of incomplete knowledge of governing parameters and failure model specification errors is expected to be important.

46 CFR 160.036-5 - Marking.

Code of Federal Regulations, 2012 CFR

2012-10-01

...: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Hand-Held Rocket-Propelled Parachute Red Flare Distress Signals § 160.036-5 Marking. (a) General. Each hand-held rocket-propelled parachute red flare distress signal shall be legibly marked or labeled as follows: (Company brand or style designation) Hand-Held Rocket...

46 CFR 160.036-5 - Marking.

Code of Federal Regulations, 2011 CFR

2011-10-01

...: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Hand-Held Rocket-Propelled Parachute Red Flare Distress Signals § 160.036-5 Marking. (a) General. Each hand-held rocket-propelled parachute red flare distress signal shall be legibly marked or labeled as follows: (Company brand or style designation) Hand-Held Rocket...

46 CFR 160.036-5 - Marking.

Code of Federal Regulations, 2013 CFR

2013-10-01

...: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Hand-Held Rocket-Propelled Parachute Red Flare Distress Signals § 160.036-5 Marking. (a) General. Each hand-held rocket-propelled parachute red flare distress signal shall be legibly marked or labeled as follows: (Company brand or style designation) Hand-Held Rocket...

46 CFR 160.036-5 - Marking.

Code of Federal Regulations, 2014 CFR

2014-10-01

...: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Hand-Held Rocket-Propelled Parachute Red Flare Distress Signals § 160.036-5 Marking. (a) General. Each hand-held rocket-propelled parachute red flare distress signal shall be legibly marked or labeled as follows: (Company brand or style designation) Hand-Held Rocket...

46 CFR 160.036-5 - Marking.

Code of Federal Regulations, 2010 CFR

2010-10-01

...: SPECIFICATIONS AND APPROVAL LIFESAVING EQUIPMENT Hand-Held Rocket-Propelled Parachute Red Flare Distress Signals § 160.036-5 Marking. (a) General. Each hand-held rocket-propelled parachute red flare distress signal shall be legibly marked or labeled as follows: (Company brand or style designation) Hand-Held Rocket...

Are executive function and impulsivity antipodes? A conceptual reconstruction with special reference to addiction

PubMed Central

Bickel, Warren K.; Jarmolowicz, David P.; Mueller, E. Terry; Gatchalian, Kirstin M.; McClure, Samuel M.

2014-01-01

Rationale Although there is considerable interest in how either executive function (EF) or impulsivity relate to addiction, there is little apparent overlap between these research areas. Objectives The present paper aims to determine if components of these two constructs are conceptual antipodes—widely separated on a shared continuum. Methods EFs and impulsivities were compared and contrasted. Specifically, the definitions of the components of EF and impulsivity, the methods used to measure the various components, the populations of drug users that show deficits in these components, and the neural substrates of these components were compared and contrasted. Results Each component of impulsivity had an antipode in EF. EF, however, covered a wider range of phenomena, including compulsivity. Conclusions Impulsivity functions as an antipode of certain components of EF. Recognition of the relationship between EF and impulsivity may inform the scientific inquiry of behavioral problems such as addiction. Other theoretical implications are discussed. PMID:22441659

Impulse noise generator--design and operation.

PubMed

Brinkmann, H

1991-01-01

In the seventies PFANDER (Pfander, 1975) proposed a screening test with an impulse noise simulator to check the particular responsivity of soldiers on vulnerability of the inner ear concerning the impulse noise-induced hearing loss. According to a system developed at the University of Oldenburg (Germany) (Klug & Radek, 1987), we have constructed an impulse noise generator designed for our specific requirements that will be presented. The simulator consists of an electrical ignited impulse noise spark gap which is supplied by a 3.5 kV high voltage source. At a distance of 1.10 m from the center of the impulse noise spark gap a peak pressure level of 155 dB with a C-Duration (Pfander, 1975) of .2 msec and with the main energy in the frequency range from 1 kHz to 2 kHz was good reproducible. It would be preferable to shift the impulse noise spectrum to lower frequencies but experimental effort has failed so far.

Reconsidering the link between impulsivity and suicidal behavior.

PubMed

Anestis, Michael D; Soberay, Kelly A; Gutierrez, Peter M; Hernández, Theresa D; Joiner, Thomas E

2014-11-01

It is widely accepted that suicidal behavior often occurs with little planning. We propose, however, that suicidal behavior is rarely if ever impulsive-that it is too frightening and physically distressing to engage in without forethought-and that suicidal behavior in impulsive individuals is accounted for by painful and fearsome behaviors capable of enhancing their capacity for suicide. We conducted a meta-analysis of the association between trait impulsivity and suicidal behavior and a critical review of research considering the impulsiveness of specific suicide attempts. Meta-analytic results suggest the relationship between trait impulsivity and suicidal behavior is small. Furthermore, studies examining a mediating role of painful and provocative behaviors have uniformly supported our model. Results from our review suggest that researchers have been unable to adequately measure impulsivity of attempts and that measures sensitive to episodic planning must be developed to further our understanding of this phenomenon. © 2014 by the Society for Personality and Social Psychology, Inc.

Nuclear Thermal Rocket/Vehicle Design Options for Future NASA Missions to the Moon and Mars

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Corban, Robert R.; Mcguire, Melissa L.; Beke, Erik G.

1995-01-01

The nuclear thermal rocket (NTR) provides a unique propulsion capability to planners/designers of future human exploration missions to the Moon and Mars. In addition to its high specific impulse (approximately 850-1000 s) and engine thrust-to-weight ratio (approximately 3-10), the NTR can also be configured as a 'dual mode' system capable of generating electrical power for spacecraft environmental systems, communications, and enhanced stage operations (e.g., refrigeration for long-term liquid hydrogen storage). At present the Nuclear Propulsion Office (NPO) is examining a variety of mission applications for the NTR ranging from an expendable, single-burn, trans-lunar injection (TLI) stage for NASA's First Lunar Outpost (FLO) mission to all propulsive, multiburn, NTR-powered spacecraft supporting a 'split cargo-piloted sprint' Mars mission architecture. Each application results in a particular set of requirements in areas such as the number of engines and their respective thrust levels, restart capability, fuel operating temperature and lifetime, cryofluid storage, and stage size. Two solid core NTR concepts are examined -- one based on NERVA (Nuclear Engine for Rocket Vehicle Application) derivative reactor (NDR) technology, and a second concept which utilizes a ternary carbide 'twisted ribbon' fuel form developed by the Commonwealth of Independent States (CIS). The NDR and CIS concepts have an established technology database involving significant nuclear testing at or near representative operating conditions. Integrated systems and mission studies indicate that clusters of two to four 15 to 25 klbf NDR or CIS engines are sufficient for most of the lunar and Mars mission scenarios currently under consideration. This paper provides descriptions and performance characteristics for the NDR and CIS concepts, summarizes NASA's First Lunar Outpost and Mars mission scenarios, and describes characteristics for representative cargo and piloted vehicles compatible with a reference 240 t-class heavy lift launch vehicle (HLLV) and smaller 120 t HLLV option. Attractive performance characteristics and high-leverage technologies associated with both the engine and stage are identified, and supporting parametric sensitivity data is provided. The potential for commonality of engine and stage components to satisfy a broad range of lunar and Mars missions is also discussed.

Space Shuttle Main Engine - The Relentless Pursuit of Improvement

NASA Technical Reports Server (NTRS)

VanHooser, Katherine P.; Bradley, Douglas P.

2011-01-01

The Space Shuttle Main Engine (SSME) is the only reusable large liquid rocket engine ever developed. The specific impulse delivered by the staged combustion cycle, substantially higher than previous rocket engines, minimized volume and weight for the integrated vehicle. The dual pre-burner configuration permitted precise mixture ratio and thrust control while the fully redundant controller and avionics provided a very high degree of system reliability and health diagnosis. The main engine controller design was the first rocket engine application to incorporate digital processing. The engine was required to operate at a high chamber pressure to minimize engine volume and weight. Power level throttling was required to minimize structural loads on the vehicle early in flight and acceleration levels on the crew late in ascent. Fatigue capability, strength, ease of assembly and disassembly, inspectability, and materials compatibility were all major considerations in achieving a fully reusable design. During the multi-decade program the design evolved substantially using a series of block upgrades. A number of materials and manufacturing challenges were encountered throughout SSME s history. Significant development was required for the final configuration of the high pressure turbopumps. Fracture control was implemented to assess life limits of critical materials and components. Survival in the hydrogen environment required assessment of hydrogen embrittlement. Instrumentation systems were a challenge due to the harsh thermal and dynamic environments within the engine. Extensive inspection procedures were developed to assess the engine components between flights. The Space Shuttle Main Engine achieved a remarkable flight performance record. All flights were successful with only one mission requiring an ascent abort condition, which still resulted in an acceptable orbit and mission. This was achieved in large part via extensive ground testing to fully characterize performance and to establish acceptable life limits. During the program over a million seconds of accumulated test and flight time was achieved. Post flight inspection and assessment was a key part of assuring proper performance of the flight hardware. By the end of the program the predicted reliability had improved by a factor of four. These unique challenges, evolution of the design, and the resulting reliability will be discussed in this paper.

Carrier rockets

NASA Astrophysics Data System (ADS)

Aleksandrov, V. A.; Vladimirov, V. V.; Dmitriev, R. D.; Osipov, S. O.

This book takes into consideration domestic and foreign developments related to launch vehicles. General information concerning launch vehicle systems is presented, taking into account details of rocket structure, basic design considerations, and a number of specific Soviet and American launch vehicles. The basic theory of reaction propulsion is discussed, giving attention to physical foundations, the various types of forces acting on a rocket in flight, basic parameters characterizing rocket motion, the effectiveness of various approaches to obtain the desired velocity, and rocket propellants. Basic questions concerning the classification of launch vehicles are considered along with construction and design considerations, aspects of vehicle control, reliability, construction technology, and details of structural design. Attention is also given to details of rocket motor design, the basic systems of the carrier rocket, and questions of carrier rocket development.

Diet-induced impulsivity: Effects of a high-fat and a high-sugar diet on impulsive choice in rats

PubMed Central

Pirkle, Jesseca R. A.; Kirkpatrick, Kimberly

2017-01-01

Impulsive choice is a common charactertistic among individuals with gambling problems, obesity, and substance abuse issues. Impulsive choice has been classified as a trans-disease process, and understanding the etiology of trait impulsivity could help to understand how diseases and disorders related to impulsive choice are manifested. The Western diet is a possible catalyst of impulsive choice as individuals who are obese and who eat diets high in fat and sugar are typically more impulsive. However, such correlational evidence is unable to discern the direction and causal nature of the relationship. The present study sought to determine how diet may directly contribute to impulsive choice. After 8 weeks of dietary exposure (high-fat, high-sugar, chow), the rats were tested on an impulsive choice task, which presented choices between a smaller-sooner reward (SS) and a larger-later reward (LL). Then, the rats were transferred to a chow diet and retested on the impulsive choice task. The high-sugar and high-fat groups made significantly more impulsive choices than the chow group. Both groups became more self-controlled when they were off the diet, but there were some residual effects of the diet on choice behavior. These results suggest that diet, specifically one high in processed fat or sugar, induces impulsive choice. This diet-induced impulsivity could be a precursor to other disorders that are characterized by impulsivity, such as diet-induced obesity, and could offer potential understanding of the trans-disease nature of impulsive choice. PMID:28662133

Diet-induced impulsivity: Effects of a high-fat and a high-sugar diet on impulsive choice in rats.

PubMed

Steele, Catherine C; Pirkle, Jesseca R A; Kirkpatrick, Kimberly

2017-01-01

Impulsive choice is a common charactertistic among individuals with gambling problems, obesity, and substance abuse issues. Impulsive choice has been classified as a trans-disease process, and understanding the etiology of trait impulsivity could help to understand how diseases and disorders related to impulsive choice are manifested. The Western diet is a possible catalyst of impulsive choice as individuals who are obese and who eat diets high in fat and sugar are typically more impulsive. However, such correlational evidence is unable to discern the direction and causal nature of the relationship. The present study sought to determine how diet may directly contribute to impulsive choice. After 8 weeks of dietary exposure (high-fat, high-sugar, chow), the rats were tested on an impulsive choice task, which presented choices between a smaller-sooner reward (SS) and a larger-later reward (LL). Then, the rats were transferred to a chow diet and retested on the impulsive choice task. The high-sugar and high-fat groups made significantly more impulsive choices than the chow group. Both groups became more self-controlled when they were off the diet, but there were some residual effects of the diet on choice behavior. These results suggest that diet, specifically one high in processed fat or sugar, induces impulsive choice. This diet-induced impulsivity could be a precursor to other disorders that are characterized by impulsivity, such as diet-induced obesity, and could offer potential understanding of the trans-disease nature of impulsive choice.

Impulsivity and the Modular Organization of Resting-State Neural Networks

PubMed Central

Davis, F. Caroline; Knodt, Annchen R.; Sporns, Olaf; Lahey, Benjamin B.; Zald, David H.; Brigidi, Bart D.; Hariri, Ahmad R.

2013-01-01

Impulsivity is a complex trait associated with a range of maladaptive behaviors, including many forms of psychopathology. Previous research has implicated multiple neural circuits and neurotransmitter systems in impulsive behavior, but the relationship between impulsivity and organization of whole-brain networks has not yet been explored. Using graph theory analyses, we characterized the relationship between impulsivity and the functional segregation (“modularity”) of the whole-brain network architecture derived from resting-state functional magnetic resonance imaging (fMRI) data. These analyses revealed remarkable differences in network organization across the impulsivity spectrum. Specifically, in highly impulsive individuals, regulatory structures including medial and lateral regions of the prefrontal cortex were isolated from subcortical structures associated with appetitive drive, whereas these brain areas clustered together within the same module in less impulsive individuals. Further exploration of the modular organization of whole-brain networks revealed novel shifts in the functional connectivity between visual, sensorimotor, cortical, and subcortical structures across the impulsivity spectrum. The current findings highlight the utility of graph theory analyses of resting-state fMRI data in furthering our understanding of the neurobiological architecture of complex behaviors. PMID:22645253

Posttraumatic Stress Disorder's Underlying Dimensions and Their Relation With Impulsivity Facets.

PubMed

Contractor, Ateka A; Armour, Cherie; Forbes, David; Elhai, Jon D

2016-01-01

Research indicates a significant relationship between posttraumatic stress disorder (PTSD) and impulsivity (Kotler, Julian, Efront, and Amir, J Nerv Ment Dis 189:162-167, 2001; Ledgerwood and Petry, J Trauma Stress 19:411-416, 2006). The present study assessed relations between PTSD symptom clusters and impulsivity subscales in an effort to assess the specific impulsivity component most related to PTSD's alterations in arousal/reactivity and alterations in mood/cognitions symptoms. In the current study, the PTSD Checklist for Diagnostic and Statistical Manual of Mental Disorders, 5th Edition, and the UPPS Impulsivity Scale were administered to a sample of 412 nonclinical subjects with a trauma history. Results indicated that PTSD's alterations in arousal/reactivity and mood/cognition factors were most related to impulsivity's sensation-seeking tendency compared with other impulsivity components. Results highlight the importance of assessing and addressing (1) sensation-seeking tendencies and (2) urges to act impulsively when experiencing negative affect in trauma treatment. Furthermore, it is possible that sensation-seeking tendencies are primarily driving the comorbidity between PTSD and certain impulsive behaviors.

Childhood impulsive behavior and problem gambling by adulthood: A 30-year prospective community-based study

PubMed Central

Shenassa, Edmond D.; Paradis, Angela D.; Dolan, Sara L.; Wilhelm, Charlotte S.; Buka, Stephen L.

2015-01-01

Aims Problem gambling can create major financial, emotional and sometimes criminal problems for an individual. This study prospectively investigated the association between impulsive behavior at age 7 and the development of lifetime problem gambling by adulthood. We also examined the specificity of any observed association between impulsive behaviors and problem gambling by conducting parallel analyses examining the link between respondents’ shy/depressed behavior in childhood and later problem gambling. Design, setting and participants Cohort study of 958 offspring of mothers enrolled in the Collaborative Perinatal Project who participated in an adult follow-up study at a mean age of 39 years. Measurements Multivariable logistic regression models were fit to determine associations between psychologist-rated impulsive and shy/depressed behaviors at age 7 and lifetime self-reported gambling as measured by the South Oaks Gambling Screen administered during the adult follow-up study. Findings Children who exhibited impulsive behaviors at age 7, compared to their non-impulsive counterparts, were 3.09 (95% confidence interval: 1.40–6.82) times as likely to report problem gambling years later. In contrast, we did not find a significant association between childhood shy/depressed behavior and problem gambling by adulthood in adjusted analyses. Conclusions Impulsive behaviors at age 7 are a specific and significant risk factor for later problem gambling.. PMID:21752146

Diagnosis and treatment of impulse control disorders in patients with movement disorders.

PubMed

Mestre, Tiago A; Strafella, Antonio P; Thomsen, Teri; Voon, Valerie; Miyasaki, Janis

2013-05-01

Impulse control disorders are a psychiatric condition characterized by the failure to resist an impulsive act or behavior that may be harmful to self or others. In movement disorders, impulse control disorders are associated with dopaminergic treatment, notably dopamine agonists (DAs). Impulse control disorders have been studied extensively in Parkinson's disease, but are also recognized in restless leg syndrome and atypical Parkinsonian syndromes. Epidemiological studies suggest younger age, male sex, greater novelty seeking, impulsivity, depression and premorbid impulse control disorders as the most consistent risk factors. Such patients may warrant special monitoring after starting treatment with a DA. Various individual screening tools are available for people without Parkinson's disease. The Questionnaire for Impulsive-Compulsive Disorders in Parkinson's Disease has been developed specifically for Parkinson's disease. The best treatment for impulse control disorders is prevention. However, after the development of impulse control disorders, the mainstay intervention is to reduce or discontinue the offending anti-Parkinsonian medication. In refractory cases, other pharmacological interventions are available, including neuroleptics, antiepileptics, amantadine, antiandrogens, lithium and opioid antagonists. Unfortunately, their use is only supported by case reports, small case series or open-label clinical studies. Prospective, controlled studies are warranted. Ongoing investigations include naltrexone and nicotine.

Study of solid rocket motors for a space shuttle booster. Volume 1: Executive summary

NASA Technical Reports Server (NTRS)

1972-01-01

An analysis of the solid propellant rocket engines for use with the space shuttle booster was conducted. A definition of the specific solid propellant rocket engine stage designs, development program requirements, production requirements, launch requirements, and cost data for each program phase were developed.

Some neuroanatomical insights to impulsive aggression in schizophrenia.

PubMed

Leclerc, Marcel P; Regenbogen, Christina; Hamilton, Roy H; Habel, Ute

2018-06-13

Patients with schizophrenia are at increased risk of engaging in violence towards others, compared to both the general population and most other patient groups. We have here explored the role of cortico-limbic impairments in schizophrenia, and have considered these brain regions specifically within the framework of a popular neuroanatomical model of impulsive aggression. In line with this model, evidence in patients with aggressive schizophrenia implicated structural deficits associated with impaired decision-making, emotional control and evaluation, and social information processing, especially in the orbitofrontal and ventrolateral prefrontal cortex. Given the pivotal role of the orbitofrontal and ventrolateral cortex in emotion control and evaluation, structural deficits may result in inappropriate use of socially relevant information and improper recognition of impulses that are in need for regulation. Furthermore, we have extended the original model and incorporated the striatum, important for the generation of aggressive impulses, as well as the hippocampus, a region critical for decision-making, into the model. Lastly, we discuss the question whether structural impairments are specific to aggressive schizophrenia. Our results suggest, that similar findings can be observed in other aggressive patient populations, making the observed impairments non-specific to aggressive schizophrenia. This points towards a shared condition, across pathologies, a potential common denominator being impulsive aggression. Copyright © 2018. Published by Elsevier B.V.

Limbic striatal dopamine D2/3 receptor availability is associated with non-planning impulsivity in healthy adults after exclusion of potential dissimulators.

PubMed

Reeves, Suzanne J; Polling, Catherine; Stokes, Paul R A; Lappin, Julia M; Shotbolt, Paul P; Mehta, Mitul A; Howes, Oliver D; Egerton, Alice

2012-04-30

Positron emission tomography (PET) studies have reported an association between reduced striatal dopamine D2/3 receptor availability and higher scores on self-report measures of trait impulsivity in healthy adults. However, impulsivity is a multi-faceted construct, and it is unclear which aspect(s) of impulsivity might be driving these associations. The current study aimed to investigate the relationship between limbic (ventral) striatal D2/3 receptor availability and individual components of impulsivity (attentional, motor and non-planning) using the Barratt Impulsiveness Scale (BIS-11) and [(11)C]raclopride PET in 23 healthy volunteers. A partial correlational analysis showed a significant association between non-planning impulsiveness (lack of forethought or 'futuring') and limbic D2/3 receptor availability, which was only apparent after the exclusion of potential dissimulators (indexed by high scores on impression management). Our findings suggest that non-planning impulsiveness is associated with individual variation in limbic striatal D2/3 receptor availability and that different facets of impulsivity may have specific neurochemical correlates. Future studies that combine D2/3 receptor imaging with behavioral measures of impulsivity are required to further elucidate the precise relationship between individual components of trait impulsivity and brain dopaminergic function. Copyright © 2012 Elsevier Ireland Ltd. All rights reserved.

Two Temperature Modeling and Experimental Measurements of Laser Sustained Hydrogen Plasmas

DTIC Science & Technology

1993-05-01

4 1.3 Theoretical Background .................................................................. 7 1.4...typically produce low specific impulses with an upper limit of approximately 450 seconds. The theoretical chamber temperature in such a system can be as...systems are theoretically capable of producing moderate thrusts (> 1 kN) with specific impulses in excess of 1000 seconds for 10 MW input power. This

Relationships between impulsivity, anxiety, and risk-taking and neural correlates of attention in adolescents

PubMed Central

Elsey, James W. B.; Crowley, Michael J.; Mencl, W. Einar; Lacadie, Cheryl M.; Mayes, Linda C.; Potenza, Marc N.

2016-01-01

Although impulsivity, anxiety, and risk-taking may relate to attentional processes, little research has directly investigated how each may be associated with specific facets of attentional processes and their underlying neural correlates. Nineteen adolescents performed an fMRI task involving simple, selective and divided attention. Out-of-scanner-assessed impulsivity, anxiety and risk-taking scores were not correlated with each other and showed task-phase-specific patterns of association. Results are discussed in light of research and theory suggesting a relationship between these domains and attention and may serve to focus future research aiming to understand these relationships. PMID:27135550

Self-reported impulsivity, but not behavioral choice or response impulsivity, partially mediates the effect of stress on drinking behavior

PubMed Central

HAMILTON, KRISTEN R.; ANSELL, EMILY B.; REYNOLDS, BRADY; POTENZA, MARC N.; SINHA, RAJITA

2013-01-01

Stress and impulsivity contribute to alcohol use, and stress may also act via impulsivity to increase drinking behavior. Impulsivity represents a multi-faceted construct and self-report and behavioral assessments may effectively capture distinct clinically relevant factors. The present research investigated whether aspects of impulsivity mediate the effect of stress on alcohol use. A community-based sample of 192 men and women was assessed on measures of cumulative stress, alcohol use, self-reported impulsivity, and behavioral choice and response impulsivity. Data were analyzed using regression and bootstrapping techniques to estimate indirect effects of stress on drinking via impulsivity. Cumulative adversity exhibited both direct effects and indirect effects (via self-reported impulsivity) on drinking behavior. Additional models examining specific types of stress indicated direct and indirect effects of trauma and recent life events, and indirect effects of major life events and chronic stressors on drinking behavior. Overall, cumulative stress was associated with increased drinking behavior, and this effect was partially mediated by self-reported impulsivity. Self-reported impulsivity also mediated the effects of different types of stress on drinking behavior. These findings highlight the value of mediation models to examine the pathways through which different types of stress increase drinking behavior. Treatment and prevention strategies should focus on enhancing stress management and self-control. PMID:22376044

Self-reported impulsivity, but not behavioral choice or response impulsivity, partially mediates the effect of stress on drinking behavior.

PubMed

Hamilton, Kristen R; Ansell, Emily B; Reynolds, Brady; Potenza, Marc N; Sinha, Rajita

2013-01-01

Stress and impulsivity contribute to alcohol use, and stress may also act via impulsivity to increase drinking behavior. Impulsivity represents a multi-faceted construct and self-report and behavioral assessments may effectively capture distinct clinically relevant factors. The present research investigated whether aspects of impulsivity mediate the effect of stress on alcohol use. A community-based sample of 192 men and women was assessed on measures of cumulative stress, alcohol use, self-reported impulsivity, and behavioral choice and response impulsivity. Data were analyzed using regression and bootstrapping techniques to estimate indirect effects of stress on drinking via impulsivity. Cumulative adversity exhibited both direct effects and indirect effects (via self-reported impulsivity) on drinking behavior. Additional models examining specific types of stress indicated direct and indirect effects of trauma and recent life events, and indirect effects of major life events and chronic stressors on drinking behavior. Overall, cumulative stress was associated with increased drinking behavior, and this effect was partially mediated by self-reported impulsivity. Self-reported impulsivity also mediated the effects of different types of stress on drinking behavior. These findings highlight the value of mediation models to examine the pathways through which different types of stress increase drinking behavior. Treatment and prevention strategies should focus on enhancing stress management and self-control.

Study of monopropellants for electrothermal thrusters

NASA Technical Reports Server (NTRS)

Kuenzly, J. D.

1974-01-01

A 333 mN electrothermal thruster designed to use MIL-grade hydrazine was demonstrated to be suitable for operation with low freezing point monopropellants containing hydrazine azide, monomethylhydrazine, unsymmetrical-dimethylhydrazine and ammonia. The steady-state specific impulse was greater than 200 sec for all propellants. The pulsed-mode specific impulse for an azide blend exceeded 175 sec for pulse widths greater than 50 msec; propellants containing carbonaceous species delivered 175 sec pulsed-mode specific impulses for pulse widths greater than 100 msec. Longer thrust chamber residence times were required for the carbonaceous propellants; the original thruster design was modified by increasing the characteristic chamber length and screen packing density. Specific recommendations were made for the work required to design and develop flight worthy thrusters, including methods to increase propellant dispersal at injection, thruster geometry changes to reduce holding power levels and methods to initiate the rapid decomposition of the carbonaceous propellants.

Project-based introduction to aerospace engineering course: A model rocket

NASA Astrophysics Data System (ADS)

Jayaram, Sanjay; Boyer, Lawrence; George, John; Ravindra, K.; Mitchell, Kyle

2010-05-01

In this paper, a model rocket project suitable for sophomore aerospace engineering students is described. This project encompasses elements of drag estimation, thrust determination and analysis using digital data acquisition, statistical analysis of data, computer aided drafting, programming, team work and written communication skills. The student built rockets are launched in the university baseball field with the objective of carrying a specific amount of payload so that the rocket achieves a specific altitude before the parachute is deployed. During the course of the project, the students are introduced to real-world engineering practice through written report submission of their designs. Over the years, the project has proven to enhance the learning objectives, yet cost effective and has provided good outcome measures.

Penrose junction conditions extended: Impulsive waves with gyratons

NASA Astrophysics Data System (ADS)

Podolský, J.; Švarc, R.; Steinbauer, R.; Sämann, C.

2017-09-01

We generalize the classical junction conditions for constructing impulsive gravitational waves by the Penrose "cut and paste" method. Specifically, we study nonexpanding impulses which propagate in spaces of constant curvature with any value of the cosmological constant (that is, Minkowski, de Sitter, or anti-de Sitter universes) when additional off-diagonal metric components are present. Such components encode a possible angular momentum of the ultrarelativistic source of the impulsive wave—the so-called gyraton. We explicitly derive and analyze a specific transformation that relates the distributional form of the metric to a new form which is (Lipschitz) continuous. Such a transformation automatically implies an extended version of the Penrose junction conditions. It turns out that the conditions for identifying points of the background spacetime across the impulse are the same as in the original Penrose cut and paste construction, but their derivatives now directly represent the influence of the gyraton on the axial motion of test particles. Our results apply both for vacuum and nonvacuum solutions of Einstein's field equations and can also be extended to other theories of gravity.

Impulsivity and reasons for living among African American youth: a risk-protection framework of suicidal ideation.

PubMed

Salami, Temilola K; Brooks, Bianca A; Lamis, Dorian A

2015-05-15

This study aims to explore the impact of specific facets of impulsivity as measured by the UPPS Impulsive Behavior Scale (UPPS), as well as reasons for living in predicting suicidal ideation among African American college-aged students. The incremental validity of each facet of the UPPS interacting with reasons for living, a construct meant to buffer against risk for suicide, was explored in a sample of African American students (N = 130; ages 18-24). Results revealed significant interactions between reasons for living and two factors of impulsivity, (lack of) premeditation and sensation seeking. Higher levels of sensation seeking and lack of premeditation in conjunction with lower reasons for living was associated with increased suicidal ideation. Neither urgency nor (lack of) perseverance significantly interacted with reasons for living in association with suicidal ideation. These results suggest including elements of impulsivity, specifically sensation seeking and (lack of) premeditation, when screening for suicidal ideation among African American youth. Future investigations should continue to integrate factors of both risk and protection when determining risk for suicide.

Impulsivity and Reasons for Living Among African American Youth: A Risk-Protection Framework of Suicidal Ideation

PubMed Central

Salami, Temilola K.; Brooks, Bianca A.; Lamis, Dorian A.

2015-01-01

This study aims to explore the impact of specific facets of impulsivity as measured by the UPPS Impulsive Behavior Scale (UPPS), as well as reasons for living in predicting suicidal ideation among African American college-aged students. The incremental validity of each facet of the UPPS interacting with reasons for living, a construct meant to buffer against risk for suicide, was explored in a sample of African American students (N = 130; ages 18–24). Results revealed significant interactions between reasons for living and two factors of impulsivity, (lack of) premeditation and sensation seeking. Higher levels of sensation seeking and lack of premeditation in conjunction with lower reasons for living was associated with increased suicidal ideation. Neither urgency nor (lack of) perseverance significantly interacted with reasons for living in association with suicidal ideation. These results suggest including elements of impulsivity, specifically sensation seeking and (lack of) premeditation, when screening for suicidal ideation among African American youth. Future investigations should continue to integrate factors of both risk and protection when determining risk for suicide. PMID:25988310

The Minimum Impulse Thruster

NASA Technical Reports Server (NTRS)

Parker, J. Morgan; Wilson, Michael J.

2005-01-01

The Minimum Impulse Thruster (MIT) was developed to improve the state-of-the-art minimum impulse capability of hydrazine monopropellant thrusters. Specifically, a new fast response solenoid valve was developed, capable of responding to a much shorter electrical pulse width, thereby reducing the propellant flow time and the minimum impulse bit. The new valve was combined with the Aerojet MR-103, 0.2 lbf (0.9 N) thruster and put through an extensive Delta-qualification test program, resulting in a factor of 5 reduction in the minimum impulse bit, from roughly 1.1 milli-lbf-seconds (5 milliNewton seconds) to - 0.22 milli-lbf-seconds (1 mN-s). To maintain it's extensive heritage, the thruster itself was left unchanged. The Minimum Impulse Thruster provides mission and spacecraft designers new design options for precision pointing and precision translation of spacecraft.

Flight motor set 360L006 (STS-34). Volume 1: System overview

NASA Technical Reports Server (NTRS)

Garecht, Diane M.

1990-01-01

Flight motor set 360L006 was launched at approximately 11:54 a.m. Central Daylight Time (CDT) on 18 October 1989 as part of NASA space shuttle mission STS-34. As with all previous redesigned solid rocket motor launches, overall motor performance was excellent. All ballistic contract end item (CEI) specification parameters were verified with the exceptions of ignition interval and rise rates. Ignition interval and rise rates could not be verified due to the elimination of developmental flight instrumentation from fourth flight and subsequent, but the low sample rate data that were available showed nominal propulsion performance. All ballistic and mass property parameters closely matched the predicted values and were well within the required CEI specification levels that could be assessed, with the exception of the RH-motor vacuum-delivered specific impulse. It exceeds the upper-limit CEI specification due to a bias imposed on the raw data by the OPT/Taber gage measurement differences. Evaluation of the ground environment instrumentation measurements again verified thermal model analysis data and showed agreement with predicted environmental effects. No launch commit criteria thermal violations occurred. Postflight inspection again verified superior performance of the insulation, phenolics, metal parts, and seals. Postflight evaluation indicated both nozzles performed as expected during flight, although splashdown loads tore the left-hand, 45-deg actuator bracket from the nozzle. All combustion gas was contained by insulation in the field and nozzle-to-case joints. Recommendations were made concerning improved thermal modeling and measurements. The rationale for these recommendations, the disposition of all anomalies, and complete result details are contained.

Orbital transfer of large space structures with nuclear electric rockets

NASA Technical Reports Server (NTRS)

Silva, T. H.; Byers, D. C.

1980-01-01

This paper discusses the potential application of electric propulsion for orbit transfer of a large spacecraft structure from low earth orbit to geosynchronous altitude in a deployed configuration. The electric power was provided by the spacecraft nuclear reactor space power system on a shared basis during transfer operations. Factors considered with respect to system effectiveness included nuclear power source sizing, electric propulsion thruster concept, spacecraft deployment constraints, and orbital operations and safety. It is shown that the favorable total impulse capability inherent in electric propulsion provides a potential economic advantage over chemical propulsion orbit transfer vehicles by reducing the number of Space Shuttle flights in ground-to-orbit transportation requirements.

Measurement of low-frequency magnetic pulses from negative stepped leaders in rocket-triggered lightning flashes

NASA Astrophysics Data System (ADS)

Lu, Gaopeng

2017-04-01

Measurement of low-frequency magnetic pulses from negative stepped leaders in rocket-triggered lightning flashes Gaopeng Lu,1,2 Yanfeng Fan,1,3 Hongbo Zhang,1,3 Rubin Jiang,1,2 Mingyuan Liu,1,2 and Xiushu Qie,1,2 1. Key Laboratory of Middle Atmosphere and Global Environment Observation, Institute of Atmospheric Physics, Chinese Academy of Sciences, Beijing, 100029, China 2. Collaborative Innovation Center on Forecast and Evaluation of Meteorological Disasters, Nanjing University of Information Science and Technology, Nanjing, Jiangsu 210044, China 3. University of Chinese Academy of Sciences, Beijing 100049, China We report the measurement of magnetic pulses from the negative stepped leaders in positive rocket-triggered lightning flashes with the low-frequency (4 kHz to 420 kHz) magnetic sensor at two different distances (78 m and 970 m, respectively) during the SHantong Artificial Triggered Lightning Experiments (SHATLE) during summer of 2015. Different from the magnetic radiation from positive leaders as observed in the considerably more frequent cases, the impulsive signals from the negative leader sustain for a much longer time interval, while the attenuation of current pulse launched by the stepping of leader is also observed. The general pattern of magnetic pulses observed for the negative stepped leader is different from the positive counterpart. Also, the initial negative leader appears to be brighter than the positive ones, as shown by both high-speed video observation and the magnetic measurement.

On the Spur of the Moment: Intrinsic Predictors of Impulse Sports Betting.

PubMed

Hing, Nerilee; Li, En; Vitartas, Peter; Russell, Alex M T

2018-06-01

Betting on impulse, without thoughtful consideration, research or informed decision-making, may cause financial and other harms and lead to the development of gambling problems. Impulse betting undermines responsible consumption of gambling because it reflects self-regulatory failure, impaired control, unreflective decision-making and betting more than planned. In this paper we define impulse gambling and report on a study that aimed to understand more about the intrinsic characteristics of sports bettors who have a greater tendency to bet on impulse. Specifically, the study aimed to identify behavioural, psychological and socio-demographic predictors of impulse sports betting. A sample of 1816 Australian sports bettors completed an online survey that measured the proportion of their bets placed on impulse both before and during sporting events, as well as bets that were researched and planned in advance. Impulse betting was common, accounting for nearly one-half of all past-year sports bets by respondents. Over three-quarters of respondents had placed one or more impulse bets in the last year and one in seven respondents had made all of their sports bets on impulse. More impulsive sports bettors were characterised as having higher trait impulsiveness, higher problem gambling severity, more frequent sports betting and a shorter history of sports betting. They favoured betting on in-match contingencies instead of overall match outcomes. While health promotion strategies are needed to discourage impulse betting, research into contextual factors that arouse urges to bet would also provide direction for harm minimisation measures that help consumers to resist impulsive betting decisions.

Affordable Development and Demonstration of a Small NTR Engine and Stage: A Preliminary NASA, DOE, and Industry Assessment

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Sefcik, Robert J.; Fittje, James E.; McCurdy, David R.; Qualls, Arthur L.; Schnitzler, Bruce G.; Werner, James E.; Weitzberg, Abraham; Joyner, Claude R.

2015-01-01

The Nuclear Thermal Rocket (NTR) represents the next evolutionary step in cryogenic liquid rocket engines. Deriving its energy from fission of uranium-235 atoms contained within fuel elements that comprise the engine's reactor core, the NTR can generate high thrust at a specific impulse of approx. 900 seconds or more - twice that of today's best chemical rockets. In FY'11, as part of the AISP project, NASA proposed a Nuclear Thermal Propulsion (NTP) effort that envisioned two key activities - "Foundational Technology Development" followed by system-level "Technology Demonstrations". Five near-term NTP activities identified for Foundational Technology Development became the basis for the NCPS project started in FY'12 and funded by NASA's AES program. During Phase 1 (FY'12-14), the NCPS project was focused on (1) Recapturing fuel processing techniques and fabricating partial length "heritage" fuel elements for the two candidate fuel forms identified by NASA and the DOE - NERVA graphite "composite" and the uranium dioxide (UO2) in tungsten "cermet". The Phase 1 effort also included: (2) Engine Conceptual Design; (3) Mission Analysis and Requirements Definition; (4) Identification of Affordable Options for Ground Testing; and (5) Formulation of an Affordable and Sustainable NTP Development Strategy. During FY'14, a preliminary plan for DDT&E was outlined by GRC, the DOE and industry for NASA HQ that involved significant system-level demonstration projects that included GTD tests at the NNSS, followed by a FTD mission. To reduce development costs, the GTD and FTD tests use a small, low thrust (approx. 7.5 or 16.5 klbf) engine. Both engines use graphite composite fuel and a "common" fuel element design that is scalable to higher thrust (approx. 25 klbf) engines by increasing the number of elements in a larger diameter core that can produce greater thermal power output. To keep the FTD mission cost down, a simple "1-burn" lunar flyby mission was considered along with maximizing the use of existing and flight proven liquid rocket and stage hardware (e.g., from the RL10-B2 engine and Delta Cryogenic Second Stage) to further ensure affordability. This paper provides a preliminary NASA, DOE and industry assessment of what is required - the key DDT&E activities, development options, and the associated schedule - to affordably build, ground test and fly a small NTR engine and stage within a 10-year timeframe.

Affordable Development and Demonstration of a Small NTR Engine and Stage: How Small is Big Enough?

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Sefcik, Robert J.; Fittje, James E.; McCurdy, David R.; Qualls, Arthur L.; Schnitzler, Bruce G.; Werner, James E.; Weitzberg (Abraham); Joyner, Claude R.

2015-01-01

The Nuclear Thermal Rocket (NTR) derives its energy from fission of uranium-235 atoms contained within fuel elements that comprise the engine's reactor core. It generates high thrust and has a specific impulse potential of approximately 900 seconds - a 100% increase over today's best chemical rockets. The Nuclear Thermal Propulsion (NTP) project, funded by NASA's AES program, includes five key task activities: (1) Recapture, demonstration, and validation of heritage graphite composite (GC) fuel (selected as the "Lead Fuel" option); (2) Engine Conceptual Design; (3) Operating Requirements Definition; (4) Identification of Affordable Options for Ground Testing; and (5) Formulation of an Affordable Development Strategy. During FY'14, a preliminary DDT&E plan and schedule for NTP development was outlined by GRC, DOE and industry that involved significant system-level demonstration projects that included GTD tests at the NNSS, followed by a FTD mission. To reduce cost for the GTD tests and FTD mission, small NTR engines, in either the 7.5 or 16.5 klbf thrust class, were considered. Both engine options used GC fuel and a "common" fuel element (FE) design. The small approximately 7.5 klbf "criticality-limited" engine produces approximately 157 megawatts of thermal power (MWt) and its core is configured with parallel rows of hexagonal-shaped FEs and tie tubes (TTs) with a FE to TT ratio of approximately 1:1. The larger approximately 16.5 klbf Small Nuclear Rocket Engine (SNRE), developed by LANL at the end of the Rover program, produces approximately 367 MWt and has a FE to TT ratio of approximately 2:1. Although both engines use a common 35 inch (approximately 89 cm) long FE, the SNRE's larger diameter core contains approximately 300 more FEs needed to produce an additional 210 MWt of power. To reduce the cost of the FTD mission, a simple "1-burn" lunar flyby mission was considered to reduce the LH2 propellant loading, the stage size and complexity. Use of existing and flight proven liquid rocket and stage hardware (e.g., from the RL10B-2 engine and Delta Cryogenic Second Stage) was also maximized to further aid affordability. This paper examines the pros and cons of using these two small engine options, including their potential to support future human exploration missions to the Moon, near Earth asteroids, and Mars, and recommends a preferred size. It also provides a preliminary assessment of the key activities, development options, and schedule required to affordably build, ground test and fly a small NTR engine and stage within a 10-year timeframe.

Ultrasonic hydrometer. [Specific gravity of electrolyte

DOEpatents

Swoboda, C.A.

1982-03-09

The disclosed ultrasonic hydrometer determines the specific gravity (density) of the electrolyte of a wet battery, such as a lead-acid battery. The hydrometer utilizes a transducer that when excited emits an ultrasonic impulse that traverses through the electrolyte back and forth between spaced sonic surfaces. The transducer detects the returning impulse, and means measures the time t between the initial and returning impulses. Considering the distance d between the spaced sonic surfaces and the measured time t, the sonic velocity V is calculated with the equation V = 2d/t. The hydrometer also utilizes a thermocouple to measure the electrolyte temperature. A hydrometer database correlates three variable parameters including sonic velocity in and temperature and specific gravity of the electrolyte, for temperature values between 0 and 40/sup 0/C and for specific gravity values between 1.05 and 1.30. Upon knowing two parameters (the calculated sonic velocity and the measured temperature), the third parameter (specific gravity) can be uniquely found in the database. The hydrometer utilizes a microprocessor for data storage and manipulation.

Lighter and Stiffer Materials for Use in Space Vehicles

NASA Astrophysics Data System (ADS)

Rioja, R. J.; Denzer, D. K.; Mooy, D.; Venema, G.

Materials for use in cryo-tanks need high specific strength, to reduce weight and high KIe fracture toughness, to avoid crack propagation through the thickness and prevent leakage. This paper presents the evolution of Aluminum alloys and products for use in rockets from the beginning of the space age to the Space Shuttle. The specific strength of alloys used in Juno, Saturn, Delta rockets and the external tank of the Space Shuttle are discussed. The microstructure and properties of 2195 plate and 2090 sheet products are reviewed. Finally, based on improved specific strength, use of 2099 and next generation upper wing 2055 Al-Li plate products is proposed for next generation rockets.

Modification of Impulse Generation During Pirouette Turns With Increased Rotational Demands.

PubMed

Zaferiou, Antonia M; Wilcox, Rand R; McNitt-Gray, Jill L

2016-10-01

This study determined how dancers regulated angular and linear impulse during the initiation of pirouettes of increased rotation. Skilled dancers (n = 11) performed single and double pirouette turns with each foot supported by a force plate. Linear and angular impulses generated by each leg were quantified and compared between turn types using probability-based statistical methods. As rotational demands increased, dancers increased the net angular impulse generated. The contribution of each leg to net angular impulse in both single and double pirouettes was influenced by stance configuration strategies. Dancers who generated more angular impulse with the push leg than with the turn leg initiated the turn with the center of mass positioned closer to the turn leg than did other dancers. As rotational demands increased, dancers tended to increase the horizontal reaction force magnitude at one or both feet; however, they used subject-specific mechanisms. By coordinating the generation of reaction forces between legs, changes in net horizontal impulse remained minimal, despite impulse regulation at each leg used to achieve more rotations. Knowledge gained regarding how an individual coordinates the generation of linear and angular impulse between both legs as rotational demand increased can help design tools to improve that individual's performance.

Human Lunar Mission Capabilities Using SSTO, ISRU and LOX-Augmented NTR Technologies: A Preliminary Assessment

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.

1995-01-01

The feasibility of conducting human missions to the Moon is examined assuming the use of three 'high leverage' technologies: (1) a single-stage-to-orbit (SSTO) launch vehicle, (2) 'in-situ' resource utilization (ISRU)--specifically 'lunar-derived' liquid oxygen (LUNOX), and (3) LOX-augmented nuclear thermal rocket (LANTR) propulsion. Lunar transportation system elements consisting of a LANTR-powered lunar transfer vehicle (LTV) and a chemical propulsion lunar landing/Earth return vehicle (LERV) are configured to fit within the 'compact' dimensions of the SSTO cargo bay (diameter: 4.6 m/length: 9.0 m) while satisfying an initial mass in low Earth orbit (IMLEO) limit of approximately 60 t (3 SSTO launches). Using approximately 8 t of LUNOX to 'reoxidize' the LERV for a 'direct return' flight to Earth reduces its size and mass allowing delivery to LEO on a single 20 t SSTO launch. Similarly, the LANTR engine's ability to operate at any oxygen/ hydrogen mixture ratio from 0 to 7 with high specific impulse (approximately 940 to 515 s) is exploited to reduce hydrogen tank volume, thereby improving packaging of the LANTR LTV's 'propulsion' and 'propellant modules'. Expendable and reusable, piloted and cargo missions and vehicle designs are presented along with estimates of LUNOX production required to support the different mission modes. Concluding remarks address the issue of lunar transportation system costs from the launch vehicle perspective.

Human lunar mission capabilities using SSTO, ISRU and LOX-augmented NTR technologies: A preliminary assessment

NASA Astrophysics Data System (ADS)

Borowski, Stanley K.

1995-10-01

The feasibility of conducting human missions to the Moon is examined assuming the use of three 'high leverage' technologies: (1) a single-stage-to-orbit (SSTO) launch vehicle, (2) 'in-situ' resource utilization (ISRU)--specifically 'lunar-derived' liquid oxygen (LUNOX), and (3) LOX-augmented nuclear thermal rocket (LANTR) propulsion. Lunar transportation system elements consisting of a LANTR-powered lunar transfer vehicle (LTV) and a chemical propulsion lunar landing/Earth return vehicle (LERV) are configured to fit within the 'compact' dimensions of the SSTO cargo bay (diameter: 4.6 m/length: 9.0 m) while satisfying an initial mass in low Earth orbit (IMLEO) limit of approximately 60 t (3 SSTO launches). Using approximately 8 t of LUNOX to 'reoxidize' the LERV for a 'direct return' flight to Earth reduces its size and mass allowing delivery to LEO on a single 20 t SSTO launch. Similarly, the LANTR engine's ability to operate at any oxygen/ hydrogen mixture ratio from 0 to 7 with high specific impulse (approximately 940 to 515 s) is exploited to reduce hydrogen tank volume, thereby improving packaging of the LANTR LTV's 'propulsion' and 'propellant modules'. Expendable and reusable, piloted and cargo missions and vehicle designs are presented along with estimates of LUNOX production required to support the different mission modes. Concluding remarks address the issue of lunar transportation system costs from the launch vehicle perspective.

"From Snail Mode to Rocket Ship Mode": Adolescents and Young Adults' Experiences of Returning to Work and School After Hematopoietic Cell Transplantation.

PubMed

Brauer, Eden R; Pieters, Huibrie C; Ganz, Patricia A; Landier, Wendy; Pavlish, Carol; Heilemann, MarySue V

2017-12-01

Resuming normal activities, such as work and school, is an important dimension of psychosocial recovery in cancer survivorship. Minimal data exist regarding adolescents or young adults' experiences of returning to school or work after cancer. The purpose of this study was to explore the processes of resuming work and school among adolescents and young adults after hematopoietic cell transplantation (HCT). In-depth interviews were conducted with 18 adolescents and young adults, who were 15-29 years when they underwent HCT and 6-60 months post-transplant at study enrollment. Interview transcripts were systematically analyzed using Grounded Theory methodology. Participants described the context in which they attempted to return to work or school, specific challenges they faced, and strategies they developed in these environments. Feeling left behind from their peers and their pre-diagnosis selves, participants described "rushing" back to school and work impulsively, taking on too much too quickly while facing overwhelming physical and cognitive demands. Factors motivating this sense of urgency as well as barriers to successful and sustainable reentry in these settings are also addressed. Findings are discussed in the context of important opportunities for clinical management, age-appropriate interventions, and implications for future research. A better understanding of psychosocial late effects, specifically related to school and work trajectories after cancer, is critical to survivorship care for adolescent and young adult cancer survivors.

Interplanetary missions with the GDM propulsion system

NASA Astrophysics Data System (ADS)

Kammash, T.; Emrich, W.

1998-01-01

The Gasdynamic Mirror (GDM) fusion propulsion system utilizes a magnetic mirror machine in which a hot dense plasma is confined long enough to produce fusion energy while allowing a fraction of its charged particle population to escape from one end to generate thrust. The particles escaping through the opposite end have their energy converted to electric power which can be used to sustain the system in a steady state operation. With the aid of a power flow diagram the minimum demands on energy production can be established and the propulsive capability of the system can be determined by solving an appropriate set of governing equations. We apply these results to several missions within the solar system and compute the trip time by invoking a continuous burn, acceleration/deceleration type of trajectory with constant thrust and specific impulse. Ignoring gravitational effects of the planets or the sun, and neglecting the change in the Earth's position during the flight we compute the round trip time for missions from Earth to Mars, Jupiter, and Pluto using linear distances and certain payload fractions. We find that a round trip to Mars with the GDM rocket takes about 170 days while those to Jupiter and Pluto take 494 and 1566 days respectively.

New technology innovations with potential for space applications

NASA Astrophysics Data System (ADS)

Krishen, Kumar

2008-07-01

Human exploration and development of space is being pursued by spacefaring nations to explore, use, and enable the development of space and expand the human experience there. The goals include: increasing human knowledge of nature's processes using the space environment; exploring and settling the solar system; achieving routine space travel; and enriching life on Earth through living and working in space. A crucial aspect of future space missions is the development of infrastructure to optimize safety, productivity, and costs. A major component of mission execution is operations management. NASA's International Space Station is providing extensive experience in both infrastructure and operations. In view of this, a vigorously organized approach is needed to implement successful space-, planet-, and ground-based research and operations that entails wise and efficient use of technical and human resources. Many revolutionary technologies being pursued by researchers and technologists may be vital in making space missions safe, reliable, cost-effective, and productive. These include: ionic polymer-metal composite technology; solid-state lasers; time-domain sensors and communication systems; high-temperature superconductivity; nanotechnology; variable specific impulse magneto plasma rocket; fuzzy logic; wavelet technology; and neural networks. An overview of some of these will be presented, along with their application to space missions.

Current Status of the Gasdynamic Mirror Fusion Propulsion Experiment

NASA Technical Reports Server (NTRS)

Emrich, William J., Jr.

2002-01-01

Nuclear fusion appears to be the most promising concept for producing extremely high specific impulse rocket engines. One particular fusion concept which seems to be particularly well suited for fusion propulsion applications is the gasdynamic mirror (GDM). An experimental GDM device has been constructed at the NASA Marshall Space Flight Center to provide an initial assessment of the feasibility of this type of propulsion system. An initial shakedown of the device is currently underway with initial experiments slated to occur in late 2001. This device would operate at much higher plasma densities and with much larger L/D ratios than previous mirror machines. The high L/D ratio minimizes to a large extent certain magnetic curvature effects which lead to plasma instabilities causing a loss of plasma confinement. The high plasma density results in the plasma behaving much more like a conventional fluid with a mean free path shorter than the length of the device. This characteristic helps reduce problems associated with 'loss cone' microinstabilities. The device has been constructed to allow a considerable degree of flexibility in its configuration thus permitting the experiment to grow over time without necessitating a great deal of additional fabrication.

Particle Streak Velocimetry of Supersonic Nozzle Flows

NASA Technical Reports Server (NTRS)

Willits, J. D.; Pourpoint, T. L.

2016-01-01

A novel velocimetry technique to probe the exhaust flow of a laboratory scale combustor is being developed. The technique combines the advantages of standard particle velocimetry techniques and the ultra-fast imaging capabilities of a streak camera to probe high speed flows near continuously with improved spatial and velocity resolution. This "Particle Streak Velocimetry" technique tracks laser illuminated seed particles at up to 236 picosecond temporal resolution allowing time-resolved measurement of one-dimensional flows exceeding 2000 m/s as are found in rocket nozzles and many other applications. Developmental tests with cold nitrogen have been performed to validate and troubleshoot the technique with supersonic flows of much lower velocity and without background noise due to combusting flow. Flow velocities on the order of 500 m/s have been probed with titanium dioxide particles and a continuous-wave laser diode. Single frame images containing multiple streaks are analyzed to find the average slope of all incident particles corresponding to the centerline axial flow velocity. Long term objectives for these tests are correlation of specific impulse to theoretical combustion predictions and direct comparisons between candidate green fuels and the industry standard, monomethylhydrazine, each tested under identical conditions.

Investigation of a Light Gas Helicon Plasma Source for the VASIMR Space Propulsion System

NASA Technical Reports Server (NTRS)

Squire, J. P.; Chang-Diaz, F. R.; Jacobson, V. T.; Glover, T. W.; Baity, F. W.; Carter, M. D.; Goulding, R. H.; Bengtson, R. D.; Bering, E. A., III

2003-01-01

An efficient plasma source producing a high-density (approx.10(exp 19/cu m) light gas (e.g. H, D, or He) flowing plasma with a high degree of ionization is a critical component of the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) concept. We are developing an antenna to apply ICRF power near the fundamental ion cyclotron resonance to further accelerate the plasma ions to velocities appropriate for space propulsion applications. The high degree of ionization and a low vacuum background pressure are important to eliminate the problem of radial losses due to charge exchange. We have performed parametric (e.g. gas flow, power (0.5 - 3 kW), magnetic field , frequency (25 and 50 MHz)) studies of a helicon operating with gas (H2 D2, He, N2 and Ar) injected at one end with a high magnetic mirror downstream of the antenna. We have explored operation with a cusp and a mirror field upstream. Plasma flows into a low background vacuum (<10(exp -4) torr) at velocities higher than the ion sound speed. High densities (approx. 10(exp 19/cu m) have been achieved at the location where ICRF will be applied, just downstream of the magnetic mirror.

Diagnosis and treatment of impulse control disorders in patients with movement disorders

PubMed Central

Mestre, Tiago A.; Strafella, Antonio P.; Thomsen, Teri; Voon, Valerie

2013-01-01

Impulse control disorders are a psychiatric condition characterized by the failure to resist an impulsive act or behavior that may be harmful to self or others. In movement disorders, impulse control disorders are associated with dopaminergic treatment, notably dopamine agonists (DAs). Impulse control disorders have been studied extensively in Parkinson’s disease, but are also recognized in restless leg syndrome and atypical Parkinsonian syndromes. Epidemiological studies suggest younger age, male sex, greater novelty seeking, impulsivity, depression and premorbid impulse control disorders as the most consistent risk factors. Such patients may warrant special monitoring after starting treatment with a DA. Various individual screening tools are available for people without Parkinson’s disease. The Questionnaire for Impulsive-Compulsive Disorders in Parkinson’s Disease has been developed specifically for Parkinson’s disease. The best treatment for impulse control disorders is prevention. However, after the development of impulse control disorders, the mainstay intervention is to reduce or discontinue the offending anti-Parkinsonian medication. In refractory cases, other pharmacological interventions are available, including neuroleptics, antiepileptics, amantadine, antiandrogens, lithium and opioid antagonists. Unfortunately, their use is only supported by case reports, small case series or open-label clinical studies. Prospective, controlled studies are warranted. Ongoing investigations include naltrexone and nicotine. PMID:23634190

Role of Serotonin and Dopamine System Interactions in the Neurobiology of Impulsive Aggression and its Comorbidity with other Clinical Disorders

PubMed Central

Seo, Dongju; Patrick, Christopher J.; Kennealy, Patrick J.

2008-01-01

Impulsive aggression is characterized by an inability to regulate affect as well as aggressive impulses, and is highly comorbid with other mental disorders including depression, suicidal behavior, and substance abuse. In an effort to elucidate the neurobiological underpinnings of impulsive aggression and to help account for its connections with these other disorders, this paper reviews relevant biochemical, brain imaging, and genetic studies. The review suggests that dysfunctional interactions between serotonin and dopamine systems in the prefrontal cortex may be an important mechanism underlying the link between impulsive aggression and its comorbid disorders. Specifically, serotonin hypofunction may represent a biochemical trait that predisposes individuals to impulsive aggression, with dopamine hyperfunction contributing in an additive fashion to the serotonergic deficit. The current paper proposes a modified diathesis-stress model of impulsive aggression in which the underlying biological diathesis may be deficient serotonergic function in the ventral prefrontal cortex. This underlying disposition can be manifested behaviorally as impulsive aggression towards oneself and others, and as depression under precipitating life stressors. Substance abuse associated with impulsive aggression is understood in the context of dopamine dysregulation resulting from serotonergic deficiency. Also discussed are future research directions in the neurobiology of impulsive aggression and its comorbid disorders. PMID:19802333

Impulsive Choice in Mice Lacking Paternal Expression of Grb10 Suggests Intragenomic Conflict in Behavior

PubMed Central

Dent, Claire L.; Humby, Trevor; Lewis, Katie; Ward, Andrew; Fischer-Colbrie, Reiner; Wilkinson, Lawrence S.; Wilkins, Jon F.; Isles, Anthony R.

2018-01-01

Imprinted genes are expressed from one parental allele only as a consequence of epigenetic events that take place in the mammalian germ line and are thought to have evolved through intragenomic conflict between parental alleles. We demonstrate, for the first time, oppositional effects of imprinted genes on brain and behavior. Specifically, we show that mice lacking paternal Grb10 make fewer impulsive choices, with no dissociable effects on a separate measure of impulsive action. Taken together with previous work showing that mice lacking maternal Nesp55 make more impulsive choices, this suggests that impulsive choice behavior is a substrate for the action of genomic imprinting. Moreover, the contrasting effect of these two genes suggests that impulsive choices are subject to intragenomic conflict and that maternal and paternal interests pull this behavior in opposite directions. Finally, these data may also indicate that an imbalance in expression of imprinted genes contributes to pathological conditions such as gambling and drug addiction, where impulsive behavior becomes maladaptive. PMID:29563147

Evidence for the Trait-Impulsivity Etiological Model in a Clinical Sample: Bifactor Structure and Its Relation to Impairment and Environmental Risk.

PubMed

Rodenacker, Klaas; Hautmann, Christopher; Görtz-Dorten, Anja; Döpfner, Manfred

2018-05-01

The trait-impulsivity etiological model assumes that a general factor (trait-impulsivity) underlies attention-deficit/hyperactivity disorder (ADHD), oppositional defiant disorder (ODD), and other externalizing disorders. We investigated the plausibility of this assumption by testing the factor structure of ADHD and ODD in a bifactor framework for a clinical sample of 1420 children between 6 and 18 years of age (M = 9.99, SD = 3.34; 85% male). Further, the trait-impulsivity etiological model assumes that ODD emerges only if environmental risk factors are present. Our results support the validity of the trait-impulsivity etiological model, as they confirm that ADHD and ODD share a strong general factor of disruptive behavior (DB) in this clinical sample. Furthermore, unlike the subdimensions of ADHD, we found that the specific ODD factor explained as much true score variance as the general DB factor. This suggests that a common scale of ADHD and ODD may prove to be as important as a separate ODD subscale to assess externalizing problems in school-age children. However, all other subscales of ADHD may not explain sufficient true score variance once the impact of the general DB factor has been taken into consideration. In accordance with the trait-impulsivity model, we also showed that all factors, but predominantly the general factor and specific inattention factor, predicted parent-rated impairment, and that predominantly ODD and impulsivity are predicted by environmental risk factors.

Impulsivity and Suicidality in Adolescent Inpatients.

PubMed

Auerbach, Randy P; Stewart, Jeremy G; Johnson, Sheri L

2017-01-01

Suicide is the second leading cause of death among adolescents, and impulsivity has emerged as a promising marker of risk. The present study tested whether distinct domains of impulsivity are differentially associated with suicide ideation, plans, and attempts. Adolescents (n = 381; boys = 106, girls = 275) aged 13-19 years (M = 15.62, SD = 1.41) were recruited from an acute, residential treatment program. Within 48 h of admission to the hospital, participants were administered structured clinical interviews assessing mental health disorders and suicidality. Following these interviews, participants completed self-report questionnaires assessing symptom severity and impulsivity. Consistent with past research, an exploratory factor analysis of our 90-item impulsivity instrument resulted in a three-factor solution: Pervasive Influence of Feelings, Feelings Trigger Action, and Lack of Follow-Through. Concurrent analysis of these factors confirmed hypotheses of unique associations with suicide ideation and attempts in the past month. Specifically, whereas Pervasive Influence of Feelings (i.e., tendency for emotions to shape thoughts about the self and the future) is uniquely associated with greater suicidal ideation, Feelings Trigger Action (i.e., impulsive behavioral reactivity to emotions) is uniquely associated with the occurrence of suicide attempts, even after controlling for current psychiatric diagnoses and symptoms. Exploratory gender analyses revealed that these effects were significant in female but not male adolescents. These findings provide new insight about how specific domains of impulsivity differentially increase risk for suicide ideation and attempts. Implications for early identification and prevention of youth suicide are discussed.

Impulsivity and Suicidality in Adolescent Inpatients

PubMed Central

Stewart, Jeremy G.; Johnson, Sheri L.

2016-01-01

Suicide is the second leading cause of death among adolescents, and impulsivity has emerged as a promising marker of risk. The present study tested whether distinct domains of impulsivity are differentially associated with suicide ideation, plans, and attempts. Adolescents (n = 381; boys = 106, girls = 275) aged 13–19 years (M = 15.62, SD = 1.41) were recruited from an acute, residential treatment program. Within 48 h of admission to the hospital, participants were administered structured clinical interviews assessing mental health disorders and suicidality. Following these interviews, participants completed self-report questionnaires assessing symptom severity and impulsivity. Consistent with past research, an exploratory factor analysis of our 90-item impulsivity instrument resulted in a three-factor solution: Pervasive Influence of Feelings, Feelings Trigger Action, and Lack of Follow-Through. Concurrent analysis of these factors confirmed hypotheses of unique associations with suicide ideation and attempts in the past month. Specifically, whereas Pervasive Influence of Feelings (i.e., tendency for emotions to shape thoughts about the self and the future) is uniquely associated with greater suicidal ideation, Feelings Trigger Action (i.e., impulsive behavioral reactivity to emotions) is uniquely associated with the occurrence of suicide attempts, even after controlling for current psychiatric diagnoses and symptoms. Exploratory gender analyses revealed that these effects were significant in female but not male adolescents. These findings provide new insight about how specific domains of impulsivity differentially increase risk for suicide ideation and attempts. Implications for early identification and prevention of youth suicide are discussed. PMID:27025937

Mechanisms of Individual Differences in Impulsive and Risky Choice in Rats

PubMed Central

Kirkpatrick, Kimberly; Marshall, Andrew T.; Smith, Aaron P.

2016-01-01

Individual differences in impulsive and risky choice are key risk factors for a variety of maladaptive behaviors such as drug abuse, gambling, and obesity. In our rat model, ordered individual differences are stable across choice parameters, months of testing, and span a broad spectrum, suggesting that rats, like humans, exhibit trait-level impulsive and risky choice behaviors. In addition, impulsive and risky choices are highly correlated, suggesting a degree of correlation between these two traits. An examination of the underlying cognitive mechanisms has suggested an important role for timing processes in impulsive choice. In addition, in an examination of genetic factors in impulsive choice, the Lewis rat strain emerged as a possible animal model for studying disordered impulsive choice, with this strain demonstrating deficient delay processing. Early rearing environment also affected impulsive behaviors, with rearing in an enriched environment promoting adaptable and more self-controlled choices. The combined results with impulsive choice suggest an important role for timing and reward sensitivity in moderating impulsive behaviors. Relative reward valuation also affects risky choice, with manipulation of objective reward value (relative to an alternative reference point) resulting in loss chasing behaviors that predicted overall risky choice behaviors. The combined results are discussed in relation to domain-specific versus domain-general subjective reward valuation processes and the potential neural substrates of impulsive and risky choice. PMID:27695580

Solid Propellant Microthruster Design, Fabrication, and Testing for Nanosatellites

NASA Astrophysics Data System (ADS)

Sathiyanathan, Kartheephan

This thesis describes the design, fabrication, and testing of a solid propellant microthruster (SPM), which is a two-dimensional matrix of millimeter-sized rockets each capable of delivering millinewtons of thrust and millinewton-seconds of impulse to perform fine orbit and attitude corrections. The SPM is a potential payload for nanosatellites to increase spacecraft maneuverability and is constrained by strict mass, volume, and power requirements. The dimensions of the SPM in the millimeter-scale result in a number of scaling issues that need consideration such as a low Reynolds number, high heat loss, thermal and radical quenching, and incomplete combustion. The design of the SPM, engineered to address these issues, is outlined. The SPM fabrication using low-cost commercial off-the-shelf materials and standard micromachining is presented. The selection of a suitable propellant and its customization are described. Experimental results of SPM firing to demonstrate successful ignition and sustained combustion are presented for three configurations: nozzleless, sonic nozzle, and supersonic nozzle. The SPM is tested using a ballistic pendulum thrust stand. Impulse and thrust values are calculated and presented. The performance values of the SPM are found to be consistent with existing designs.

Impulsivity-like Traits and Risky Driving Behaviors among College Students

PubMed Central

Murphy, Elaine M.; Doane, Ashley N.

2017-01-01

The present study examined the predictive effects of five impulsivity-like traits (Premeditation, Perseverance, Sensation Seeking, Negative Urgency, and Positive Urgency) on driving outcomes (driving errors, driving lapses, driving violations, cell phone driving, traffic citations, and traffic collisions). With a convenience sample of 266 college student drivers, we found that each of the impulsivity-like traits was related to multiple risky driving outcomes. Positive Urgency (tendency to act impulsively when experiencing negative affect) was the most robust predictor of risky driving outcomes. Positive Urgency is a relatively newly conceptualized impulsivity-like trait that was not examined in the driving literature previously, suggesting a strong need to further examine its role as a personality trait related to risky driving. These findings generally support the multidimensional assessment of impulsivity-like traits, and they specifically support the addition of Positive Urgency to a list of risk factors for risky driving behaviors. PMID:23428428

Impulsivity-like traits and risky driving behaviors among college students.

PubMed

Pearson, Matthew R; Murphy, Elaine M; Doane, Ashley N

2013-04-01

The present study examined the predictive effects of five impulsivity-like traits (Premeditation, Perseverance, Sensation Seeking, Negative Urgency, and Positive Urgency) on driving outcomes (driving errors, driving lapses, driving violations, cell phone driving, traffic citations, and traffic collisions). With a convenience sample of 266 college student drivers, we found that each of the impulsivity-like traits was related to multiple risky driving outcomes. Positive Urgency (tendency to act impulsively when experiencing negative affect) was the most robust predictor of risky driving outcomes. Positive Urgency is a relatively newly conceptualized impulsivity-like trait that was not examined in the driving literature previously, suggesting a strong need to further examine its role as a personality trait related to risky driving. These findings generally support the multidimensional assessment of impulsivity-like traits, and they specifically support the addition of Positive Urgency to a list of risk factors for risky driving behaviors. Copyright © 2013 Elsevier Ltd. All rights reserved.

78 FR 42430 - Revisions to the Export Administration Regulations Based on the 2012 Missile Technology Control...

Federal Register 2010, 2011, 2012, 2013, 2014

2013-07-16

.... This specific decontrol helps to clarify what is usable for rockets, missiles, or unmanned aerial... specified by 4.C. in the Annex or used in Category I rocket systems. These commodities and the related... propulsion data into a flight management system, designed or modified for rockets or missiles capable of...

The influence of impulsiveness on binge eating and problem gambling: A prospective study of gender differences in Canadian adults.

PubMed

Farstad, Sarah M; von Ranson, Kristin M; Hodgins, David C; El-Guebaly, Nady; Casey, David M; Schopflocher, Don P

2015-09-01

This study investigated the degree to which facets of impulsiveness predicted future binge eating and problem gambling, 2 theorized forms of behavioral addiction. Participants were 596 women and 406 men from 4 age cohorts randomly recruited from a Canadian province. Participants completed self-report measures of 3 facets of impulsiveness (negative urgency, sensation seeking, lack of persistence), binge-eating frequency, and problem-gambling symptoms. Impulsiveness was assessed at baseline, and assessments of binge eating and problem gambling were followed up after 3 years. Weighted data were analyzed using zero-inflated negative binomial and Poisson regression models. We found evidence of transdiagnostic and disorder-specific predictors of binge eating and problem gambling. Negative urgency emerged as a common predictor of binge eating and problem gambling among women and men. There were disorder-specific personality traits identified among men only: High lack-of-persistence scores predicted binge eating and high sensation-seeking scores predicted problem gambling. Among women, younger age predicted binge eating and older age predicted problem gambling. Thus, there are gender differences in facets of impulsiveness that longitudinally predict binge eating and problem gambling, suggesting that treatments for these behaviors should consider gender-specific personality and demographic traits in addition to the common personality trait of negative urgency. (c) 2015 APA, all rights reserved).