FY21 Rotating Detonation Rocket Engine Annual Review Presentation
Annual review presentation for the Rotating Detonation Rocket Engine project under the Game Changing Development program.
Detonation↗
Engineering topics
Propulsion
Explore propulsion reports from NASA NTRS and DOE OSTI. Search rocket engines, electric propulsion and thrusters; verify designs and performance in the source documents.
Engine Cycle Comparison for Alternative Propellant Nuclear Thermal Propulsion Engines
A modular Nuclear Thermal Propulsion (NTP) engine modeling suite was coded in Simulink to analyze various engine cycle configurations and propellants. This model was validated and used for previous studies involving alternative propellants for NTP engines. The current study compared hydrogen-based NTP (H-NTP) with ammonia-based Alternative propellant NTP (A-NTP) engines with expander and bleed cycle configurations. Based on prior work, the bleed cycle configuration was modified to yield a feasible result. The assumption that was made was that the Testing Reference Design (TRD) reactor will be used for all cases and that the difference in fuel loading fractions will be enough to offset any differences in the neutronics caused by switching the propellant. This work provided and compared propellant state points throughout the cycles as well as the key performance parameters. The conclusion was that bleed cycles for H-NTP engines are advantageous over expander cycles in terms of maximum system pressure, cycle simplicity, and potentially engine mass. Conversely, expander cycles are more advantageous for A-NTP engines according to the same scope of parameters. Future work will analyze reactor designs that are different from the TRD and use neutronics software integrated with the X-NTP model to provide higher fidelity results.
Nuclear↗
Calculation of small rocket trajectories using only rocket-to-launcher radial velocity data
Calculating small rocket trajectories using rocket to launcher radial velocity data
DOPPLER EFFECT↗
Introduction to the problem of rocket base heating, and to the behavior of liquid propellant rocket jet plumes
Rocket vehicle base heating problem, and liquid rocket propellant plume characteristics
Dahm, W. K.↗
Propellant selection for unmanned spacecraft propulsion systems. Volume 2 - Analysis of propellant sensitivity, secondary propulsion, and ground operations, phase 2 Final report
Analysis of propellant sensitivity, secondary propulsion, and ground operations for unmanned spacecraft propulsion systems
Source record↗
Applications of variable property heat- transfer and friction equations to rocket nozzle coolant passages and comparison with nuclear rocket test results
Variable property heat transfer and friction equations to rocket nozzle coolant passages and comparison with nuclear rocket test results
Taylor, M. F.↗
Synthesis of calculational methods for the design and analysis of radiation shields for nuclear rocket systems. Volume 1 - Radiation measurements in a nuclear rocket propellant tank mockup using simulated liquid hydrogen
Calculational methods for nuclear rocket radiation shield design - radiation measurements in nuclear rocket propellant tank mockup using simulated liquid hydrogen
Rowland, J.↗
Synthesis of calculational methods for the design and analysis of radiation shields for nuclear rocket systems. Volume 2 - Analysis of radiation measurements in a nuclear rocket propellant tank mockup using simulated liquid hydrogen
Calculational methods for nuclear rocket radiation shield design - analysis of radiation measurements in nuclear rocket propellant tank mockup using simulated liquid hydrogen
Disney, R. K.↗
Primary electric propulsion in the evolving automated space transportation propulsion program
Primary solar electric propulsion for automated space transportation propulsion program, considering performance and technology readiness
Edgecombe, D. S.↗
Satellite auxiliary-propulsion selection techniques. Addendum: A survey of auxiliary electric propulsion systems
A review of electric thrusters for satellite auxiliary propulsion was conducted at JPL during the past year. Comparisons of the various thrusters for attitude propulsion and east-west and north-south stationkeeping were made based upon performance, mass, power, and demonstrated life. Reliability and cost are also discussed. The method of electrical acceleration of propellant served to divide the thruster systems into two groups: electrostatic and electromagnetic. Ion and colloid thrusters fall within the electrostatically accelerated group while MPD and pulsed plasma thrusters comprise the electromagnetically accelerated group. The survey was confined to research in the United States with accent on flight and flight prototype systems.
Holcomb, L. B.↗
Longitudinal aerodynamic and propulsion characteristics of a propulsive-wing V/STOL model at high subsonic speeds
The aerodynamic and propulsion characteristics of a 1/6-scale propulsive-wing V/STOL air-powered model was investigated over the Mach number range from 0.40 to 0.96 and at angles of attack from -5 deg to 15 deg for several fan rotational speeds. Three fanduct-exit configurations were tested, including two exit areas. The model with 25-percent-thick wing had a drag-rise Mach number of 0.85, which is typical of aircraft with thinner, conventional, unswept wings.
Salters, L. B., Jr.↗
Coordinated rocket and satellite measurements of an auroral event. II - The rocket observations and analysis
A rocket-borne payload launched into an aurora and a simultaneous overpass of the Atmosphere Explorer C satellite yielded measurements of auroral optical emission rates, thermal ion and electron densities, and low-energy electron fluxes. Model calculations of the thermospheric manifestation of the aurora were performed through use of rocket-determined auroral ionization rates and satellite-determined neutral gas densities. Measured oxygen densities provided a means of assessing the quenching rate of an excited state of N2. Energy transfer from this excited state appears to be the major source of 5577-A emission. Optical emission at 6300 A cannot be explained either by electron impact on atomic oxygen or by dissociative recombination of O2(+).
Sharp, W. E.↗
Optimization of extended propulsion time nuclear-electric propulsion trajectories
This paper presents the methodology used in optimizing extended propulsion time NEP missions considering realistic thruster lifetime constraints. These missions consist of a powered spiral escape from a 700-km circular orbit at the earth, followed by a powered heliocentric transfer with an optimized coast phase, and terminating in a spiral capture phase at the target planet. This analysis is most applicable to those missions with very high energy requirements such as outer planet orbiter missions or sample return missions where the total propulsion time could greatly exceed the expected lifetime of an individual thruster. This methodology has been applied to the investigation of NEP missions to the outer planets where examples are presented of both constrained and optimized trajectories.
Sauer, C. G., Jr.↗
Study of electrical and chemical propulsion systems for auxiliary propulsion of large space systems. Volume 1: Executive summary
The objective was to determine the direction auxiliary propulsion research and development should take to best meet upcoming needs. The approach used was to define the important electrical and chemical propulsion characteristics in terms of the demands that will be imposed by future spacecraft. Comparison of these desired characteristics and capabilities with those presently available was then used to identify deficiencies.
Smith, W. W.↗
Development of sensors for ceramic components in advanced propulsion systems: Survey and evaluation of measurement techniques for temperature, strain and heat flux for ceramic components in advanced propulsion systems
The report presents the final results of Tasks 1 and 2, Development of Sensors for Ceramic Components in Advanced Propulsion Systems (NASA program NAS3-25141). During Task 1, an extensive survey was conducted of sensor concepts which have the potential for measuring surface temperature, strain and heat flux on ceramic components for advanced propulsion systems. Each sensor concept was analyzed and evaluated under Task 2; sensor concepts were then recommended for further development. For temperature measurement, both pyrometry and thermographic phosphors are recommended for measurements up to and beyond the melting point of ceramic materials. For lower temperature test programs, the thin-film techniques offer advantages in the installation of temperature sensors. Optical strain measurement techniques are recommended because they offer the possibility of being useful at very high temperature levels. Techniques for the measurement of heat flux are recommended for development based on both a surface mounted sensor and the measurement of the temperature differential across a portion of a ceramic component or metallic substrate.
Atkinson, W. H.↗
Liquid Rocket Booster (LRB) for the Space Transportation System (STS) systems study. Appendix C: Battery report for the liquid rocket booster TVC actuators
The actuators for control of engine valves and gimbals for a booster require 165 kW or more peak power at 270 volts direct current (VDC) during the 2 or 3 minutes of first stage ascent; other booster devices require much less power at 28 VDC. It is desired that a booster supply its own electrical power and satisfy redundancy requirements of the Solid Rocket Booster Shuttle, when applicable. The power of a Liquid Rocket Booster is therefore provided by two subsystems: Actuator Battery Power (270 VDC) Subsystem for the engine actuators, and Electrical Power and Distribution (28 VDC) Subsystem, to power everything else. Boosters will receive no electrical power from Orbiter, only commands and data, according to current plans. It was concluded that nine 30 volt silver-zinc batteries-in-series be used to provide the 270 volt, 37 kW average (165 kW peak).
Source record↗
Liquid Rocket Booster (LRB) for the Space Transportion System (STS) systems study. Appendix D: Trade study summary for the liquid rocket booster
Trade studies plans for a number of elements in the Liquid Rocket Booster (LRB) component of the Space Transportation System (STS) are given in viewgraph form. Some of the elements covered include: avionics/flight control; avionics architecture; thrust vector control studies; engine control electronics; liquid rocket propellants; propellant pressurization systems; recoverable spacecraft; cryogenic tanks; and spacecraft construction materials.
Source record↗
Solid rocket combustion simulator technology using the hybrid rocket for simulation
The hybrid rocket is reexamined in light of several important unanswered questions regarding its performance. The well-known heat transfer limited burning rate equation is quoted, and its limitations are pointed out. Several inconsistencies in the burning rate determination through fuel depolymerization are explicitly discussed. The resolution appears to be through the postulate of (surface) oxidative degradation of the fuel. Experiments are initiated to study the fuel degradation in mixtures of nitrogen/oxygen in the 99.9 percent/0.1 percent to 98 percent/2 percent range. The overall hybrid combustion behavior is studied in a 2 in-diameter rocket motor, where a PMMA tube is used as the fuel. The results include detailed, real-time infrared video images of the combustion zone. Space- and time-averaged images give a broad indication of the temperature reached in the gases. A brief outline is shown of future work, which will specifically concentrate on the exploration of the role of the oxidizer transport to the fuel surface, and the role of the unburned fuel that is reported to escape below the classical time-averaged boundary layer flame.
Ramohalli, Kumar↗