Browse Topic: Rocket engines

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Large Eddy Simulations of Supercritical and Transcritical Jet Flows Using Real Fluid Thermophysical Properties2020-01-11534/14/2020
In order to understand supercritical jet flows further, well resolved large eddy simulations (LES) of a n-dodecane jet mixing with surrounding nitrogen are conducted. A real fluid thermodynamic model is used to account for the fuel compressibility and variable thermophysical properties due to the solubility of ambient gas and liquid jet using the cubic Peng-Robinson equation of state (PR-EOS). A molar averaged homogeneous mixing rule is used to calculate the mixing properties. The thermodynamic model is coupled with a pressure-based solver to simulate multispecies reacting flows. The numerical model is based on a second order accurate method implemented in the open source OpenFOAM-6 software. First, to evaluate the present numerical model for sprays, 1D advection and shock tube benchmark problems at supercritical conditions are shown. Second, a cryogenic nitrogen injection with a jet velocity of 4.9 m/s into a supercritical nitrogen environment at 4 MPa and room temperature is considered to carry out a grid resolution study, and the corresponding results are evaluated against experimental data of Mayer et al. Then, to assess the effects of thermophysical property variations due to mixing of two species, a high-pressure jet of n-dodecane at transcritical and supercritical temperatures at 200 m/s into high-pressure and high-temperature nitrogen environment is studied. Detailed analysis of the species dispersion and mixing are presented for various conditions. The present LES simulations of n-dodecane jets show massive shear forces and high hydrodynamic pressure fluctuations caused by the high-speed jet. The predicted results of flow and thermophysical properties are in close agreement with the available literature data.
Ningegowda, B M, Rahantamialisoa, Faniry, Zembi, Jacopo, Pandal, Adrian, Im, Hong G., Battistoni, Michele
Facility Focus: NASA Goddard Space Flight CenterTBMG-359692/1/2020
Dr. Robert Hutchings Goddard is considered the father of modern rocket propulsion. A physicist, Goddard also had a unique genius for invention. It is in his honor that NASA's Goddard Space Flight Center in Greenbelt, MD, was established on May 1, 1959 as NASA's first space flight complex.
Effects of Engine Speed on the Performance at Extreme Vehicle Driving Conditions2019-36-02971/13/2020
Nowadays, improvements in engine fuel economy and reduction in pollutant emissions have been much discussed. Downsizing and downspeeding are methods widely used in the automotive market, used to increase internal combustion engine efficiency. Fuel consumption is commonly measured through a specific cycle that could differ between countries. Federal Test Procedure (FTP-75) for Latin America and New European Driving Cycle (NEDC) for Europe are some examples. These cycles include normal driving conditions, low vehicle load and mild environmental conditions. The style of gear shifting is one of the main factors affecting fuel efficiency and performance of motor vehicles. At extreme diving conditions, parameters like coolant and air charge temperature could reach limit durability values of the system. In order to avoid knocking and maintain structural reliability, it becomes mandatory to control the engine speed. An increase in engine speed reduces engine load requirements and thus, turbocharger loads. Lower turbocharger requirements reduce ACT and, consequently, reduce knocking. This process allows an advanced ignition timing, enabling more chamber pressure and improving fuel conversion efficiency. However, a higher engine speed also increases friction losses and combustion frequency, lowering the generated torque and tending to decrease engine efficiency. It is interesting to note parameters that affect positively and negatively at the same time, depending on the point of view. The present paper aims to analyze the performance and fuel consumption of a turbocharged engine working under extreme driving conditions, simulating a CVT transmission behavior at various speeds. The evaluated vehicle parameters are: power, coolant temperature, ACT and brake specific fuel consumption (BSFC).
Thomaz, Fabrício, Baeta, José Guilherme Coelho
Metallization for SiC SemiconductorsTBMG-3570412/1/2019
Scientists at NASA’s Glenn Research Center have developed and patented improved ohmic contacts, and a fabrication method for them, that will speed the production of silicon carbide (SiC) sensors and electronics that operate in high-temperature, harsh environments.
Phase-Change Coating Absorbs Heat from Rockets, Pipes, and BeerTBMG-3544611/1/2019
Spinoff is NASA's annual publication featuring successfully commercialized NASA technology. This commercialization has contributed to the development of products and services in the fields of health and medicine, consumer goods, transportation, public safety, computer technology, and environmental resources.
Celebrating the 50th Anniversary of Apollo 11: Getting There, To a New Era of Space ExplorationTBMG-347457/1/2019
For the first time in a generation, NASA is building a human spacecraft for deep-space missions that will usher in a new era of space exploration. A series of increasingly challenging missions awaits, and this new spacecraft will take astronauts farther than they’ve gone before, including to the Moon and Mars.
An Update on AVCP's Zero-Zero Safety System for VTOL AircraftF-0075-2019-145555/13/2019
The advent of multi-rotor, electric VTOL aircraft capable of carrying passengers and providing on-demand personal air transport promises a revolution not only in urban travel, but also in a multitude of other applications which will change the way that a broad range of people carry out their routine work activities. Many of the eVTOL design projects propose to use current ballistic parachute systems in case of any power or control failure; however, any recovery system which (i) gives a Ground Contact Velocity (GCV) of 5 to 8m/s and (ii) will not deploy properly below c. 100 to 150 feet will not be a solution that should satisfy certification requirements for carrying fare-paying passengers above a crowded urban environment. This paper updates the detailed design and function of the rocket motors and control system designed to provide a controlled landing at less than 2m/s in any emergency situation.
Sloman, Roger
LED Lighting Improves Efficiency and ImagingTBMG-343975/1/2019
Rocket engine testing requires a lot of light since all tests are filmed with high-speed cameras to monitor performance; however, those cameras have to adjust to the bright plume from a firing engine, which would black out the rest of the image. Traditionally, that light has been provided by metal halide bulbs.
A Target Cascading Method Using Model Based Simulation in Early Stage of Vehicle Development2019-01-08364/2/2019
In the early stages of vehicle development, it is important for decision makers to understand a feasible constraint region that satisfies all system level requirements. The purpose of this paper is to propose a target cascading method to solve for a feasible design region which satisfies all constraints of the system based on model based simulation. In this method, the feasible design region is explored by using both global optimization methods and active learning techniques. In optimization problems, the inverse problem for understanding feasibility for specific designs is defined and solved. To determine the objective functions of the inverse problem, an index representing the achievement level of constraints from system requirements is introduced. To predict feasible regions in the specific design space, a surrogate model of minimized values of the index is trained by using a kriging model. Training data of the surrogate model is sequentially generated based on the expected improvement function to improve accuracy of feasible regions effectively.
Shintani, Kohei, Abe, Atsuji, Yamamoto, Yasushi
A Multi-Domain Component Based Modeling Toolset for Dynamic Integrated Power and Thermal System Modeling2019-01-13853/19/2019
Design of modern aircraft relies heavily on modeling and simulation for reducing cost and improving performance. However, the complexity of aircraft architectures requires accurate modeling of dynamic components across many subsystems. Integrated power and thermal modeling necessitates dynamic simulations of liquid, air, and two-phase fluids within vapor cycle system components, air cycle machine and propulsion components, hydraulic components, and more while heat generation of many on-board electrical components must also be precisely calculated as well. Integration of these highly complex subsystems may result in simulations which are too computationally expensive for quickly modeling extensive variations of aircraft architecture, or will require simulations with reduced accuracy in order to provide computationally inexpensive models. As such, a need for software toolsets with the ability to model complex aircraft architectures with accurate calculations while maintaining high computational speeds is apparent. This paper details the development of the ATTMOSphere toolset which enables modeling of electrical, mechanical, thermal, fluid flow and heat transfer across a range of components applicable to integrated power and thermal systems. Graphical user interfaces provide user-friendly parameterization of components, as well as sizing of many of the available components. All ATTMOSphere components operate within universal mechanical, thermal, electrical, and fluid domains allowing for seamless integration of components across many architectures, providing end-users with the ability to simultaneously model vapor cycle systems, air cycle systems, pumped refrigeration systems, and other power and thermal systems along with their interactions with parallel subsystems. This paper provides details of the components developed in ATTMOSphere along with examples of user interfaces and design codes. Demonstration models are presented to illustrate the integrated dynamic analysis capability of the toolset.
McCarthy, Patrick Thomas, McCarthy, Kevin, Hasan, Maher, Boyd, Michelle, Chang, Michael, Walters, Eric, Niedbalski, Nicholas
Interoperable Intelligent Controllers for Process Management and Control NetworksTBMG-339293/1/2019
NASA Johnson Space Center developed reprogrammable and interchangeable electronic controllers that can attach to a system or subsystem wirelessly or through plug-and-play capability. Originally designed to work with rocket engines, this technology can control different systems and subsystems. This smart controller recognizes which system it is communicating with once connected to the network and loads the appropriate application to perform the required function in the system. The device enables a common set of spares and can talk to other devices of its kind by relaying information and instructions to other controllers. A prototype can be easily developed using low-cost solutions such as Raspberry Pi to demonstrate functionality.
Integrated Multiphysics and Advanced DiagnosticsTBMG-328329/1/2018
The objective of this effort was to design, fabricate, integrate, and fly a nuclear thermal rocket without having to build massive ground test facilities. Furthermore, this nuclear rocket would be inspected internally without having to disassemble it. Another aim here was to predict the reliability for a 15-to-20-year lifetime for a surface nuclear power plant essential for supplying power to explorers on the Moon or Mars, all without having to build and test a full-scale reactor. If the performance of these systems could be predicted, one could establish the impact of manufacturing tolerances on the operations of these systems.
High Strength Elevated Temperature Bolting PracticeARP700B (Historical)8/23/2018
This SAE Aerospace Recommended Practice (ARP) provides general information on the design and installation of threaded fasteners in high strength and high temperature applications in propulsion systems. Some of the more common definitions of fastener terminology are also provided.
E-25 General Standards for Aerospace and Propulsion Systems
A Novel Zero-Zero Safety System for VTOL AircraftF-0074-2018-127405/14/2018
ABSTRACT The advent of multi-rotor, electric VTOL aircraft capable of carrying passengers and providing on-demand personal air transport promises a revolution not only in urban travel, but also in a multitude of other applications which will change the way that a broad range of people carry out their routine work activities. Many of the eVTOL design projects propose to use current ballistic parachute systems in case of any power or control failure; however, any recovery system which (i) gives a Ground Contact Velocity (GCV) of 5 to 7m/s and (ii) will not deploy properly below c. 100 to 150 feet will not be a solution that can satisfy certification requirements for carrying fare-paying passengers above a crowded urban environment. Active VTOL Crash Prevention Ltd (AVCP) is developing a system combining a rapid opening parachute and rocket motors designed to provide a controlled landing at less than 2m/s in any emergency situation.
Sloman, Roger, Scott, Simon
Freeform Channel Close-Out for High-Performance Heat ExchangersTBMG-288405/1/2018
Large-scale liquid rocket engines with regeneratively cooled nozzles will enable reliable and reduced-cost access to space. The coolant that circulates through the internal channels within the nozzle helps maintain adequate wall temperatures to prevent failure, and is contained under high pressure. It has been a challenge to affordably manufacture the intricate nozzle channels and close-out of these channels. As such, NASA has leveraged cladding and additive manufacturing technology to build the nozzle liner and outer jacket that closes out the channels within, and contains the high-pressure coolant fluid. The new capability reduces the time to fabricate the nozzle and allows for real-time inspection during the build. It has been demonstrated on a series of different alloys, hot-fire testing is nearly complete, and micrograph examination has verified that the bonds are reliable with little deformation to the coolant channels. The technology provides new opportunities to manufacture nozzles and chambers and reduce costs.
Exploring the Potential of Combustion on Titan01-11-01-00024/7/2018
Significant attention has been focused on Mars due to its relative proximity and possibility of sustaining human life. However, its lack of in-situ sources of energy presents a challenge to generate needed energy on the surface. Comparatively, Titan has a nearly endless source of fuel in its atmosphere and lakes, but both are lacking in regards to their oxidizing capacity. The finding of a possible underground liquid ammonia-water lake on Titan suggests that oxygen might actually be within reach. This effort provides the first theoretical study involving a primary energy generation system on Titan using the atmosphere as a fuel and underground water as the source for the oxygen via electrolysis from wind generated electricity. Thermodynamic calculations and use of chemical kinetics in a zero-dimensional Homogeneous Charge Compression Ignition (HCCI) engine model demonstrate that is indeed possible to operate an internal combustion engine on the surface of Titan while providing heat for terraforming and human activities. Subsequent terraforming estimates illustrate that while the potential for energy and heat exists, the amount of needed hardware is largely impractical. However, the findings may stimulate further curiosity by others to look towards outer space and imagine what might be possible.
Depcik, Christopher
Non-Catalytic Ignition System for High-Performance Advanced Monopropellant ThrustersTBMG-287324/1/2018
Anon-catalytic pyrotechnic ignition system was developed for advanced green monopropellant systems operating with hydroxylammonium nitrate (HAN)-based monopropellant AF-M315E. This technology will provide increased performance and new operating regimes for future NASA missions that employ green propulsion systems. The non-catalytic ignition system is a critical component for future high-thrust, in-space rocket engines that utilize AF-M315E, offering longer thruster life and faster startup than catalytic engines. This work demonstrated the pyrotechnic ignition system in proof-of-concept hot-fire tests showing AF-M315E combustion in a workhorse thruster in a laboratory environment.
Efficient Packaging Process for One-Piece Deployable Thin MembraneTBMG-285403/1/2018
NASA’s Marshall Space Flight Center is building a small CubeSat that uses an 85-m2 solar sail deployed from a central location to capture the push of photons from the Sun as its propulsion source. To integrate this sail into the CubeSat, NASA inventors have developed a new method for folding and packaging the thin membrane.
Cryogenic and Non-Cryogenic Liquid Level InstrumentTBMG-281851/1/2018
NASA’s Marshall Space Flight Center has developed a unique prototype for measuring the liquid level in a tank, employing a novel process. The technology can operate in a wide range of environments, including high and low temperatures and pressures, and is simpler and less expensive than other optical sensing techniques. The instrument also provides far greater accuracy and faster results in cryogenic conditions than typical cryogenic liquid metering methods. It is ideal for cryogenic and non-cryogenic ground tank metering applications, and zero-gravity systems that include stratification or settling techniques.
Sequential/Simultaneous Multi-Metallized Nanocomposites (S2M2N)TBMG-2798712/1/2017
NASA Langley Research Center has developed a method to create Sequential/Simultaneous Multi-Metallized Nanocomposites (S2M2N) via supercritical fluid (SCF) sequential or simultaneous multi-metal infusion. The SCF infusion process provides deep impregnation of metal nanoparticles into a variety of materials, including those with challenging topographies and complex structures. The resulting multi-metallized nanocomposites can possess high electrical conductivity, permittivity, permeability, wear resistance/anti-penetrant, and radiation shielding along with high toughness.
Novel Design of an Orifice Control ElementTBMG-2800612/1/2017
An orifice element is commonly used in liquid rocket engine test facilities either as a flow metering device, a damper for acoustic resonance, or to provide a large reduction in pressure over a very small distance in the piping system. The orifice as a device is largely effective in stepping down pressure; however, it is also susceptible to a wake-vortex instability that generates pressure fluctuations that propagate downstream and interact with other elements of the test facility, resulting in structural vibrations. Exacerbating the situation in cryogenic test facilities is the possibility of the formation of vapor clouds when the pressure in the wake falls below the vapor pressure, leading to cavitation. Cavitation has the potential for highamplitude fluctuations that can cause catastrophic damage to a facility.
Effect of High Frequency Acoustic Field on Atomization Behavior of Ethanol and Kerosene2017-01-231810/8/2017
Combustion instability often occurs inside the combustion chamber of aero engine. Fuel atomization and evaporation, one of the controlling processes of combustion rate, is an important mechanism of the combustion instability. To tackle combustion instability, it challenges a deep understanding of the underlying mechanism of fuel atomization and evaporation. In this paper, acoustic field was established to simulate the pressure oscillation. Transient spray images of ethanol and kerosene were recorded using high-speed camera. The obtained images were processed by MATLAB to extract and analyze the related data. Spatial fuel atomization characteristics was analytically examined by multi-threshold image method to analyze the effect of the high frequency acoustic field on the fuel break-up and disintegration. The results show that the half spray cone angle on the side with speaker is suppressed by the presence of the imposed acoustic field compared with the case without speaker. Statistically, the half spray angle of kerosene with right speaker under the acoustic frequency of 9 kHz is 18.97% larger than that with left speaker. For ethanol, the difference is 11.90%. Also, it turns out the frequency of the acoustic field influences the spray angle variation during the injection process. There is a decline of the spray cone angle at the early stage of the injection. However, the watersheds for the transformation were not the same for kerosene and ethanol.
Jia, Xiaoxu, Huang, Zhong, Ju, Dehao, Huang, Zhen, Lu, Xing-cai
Numerical Analysis on the Injection and Atomization Characteristics of Diesel Surrogates at Engine Conditions2017-01-230610/8/2017
Recently, the shortage of fossil resources contributes to strict regulations of environmental protection. The research on the high efficiency and low emission of engines becomes an important direction all over the world. Technologies like high injection pressure, high levels of supercharging and higher levels of back pressure have come into application. Increasing the injection pressure and average cylinder pressure results in that parts of the spray can experience transcritical and supercritical regimes. In this paper, we established a surrogate fuel composed of n-Hexadecane, HMN and 1-Metylnaphthalene, to analyze the injection and atomization of diesel surrogate fuel with large eddy simulation (LES) in a cubic calculation region with high temperature and high pressure environment. The injection pressure was fixed to 150MPa, and the 900-K temperature and the 6-MPa pressure represented the ambient condition in constant volume vessel which is supercritical with respect to No.2 diesel. Analyses of No.2 diesel and the surrogate fuel were performed to compare the characteristics of injection and atomization. To illustrate the differences, liquid penetration, jet penetration and spray cone angle of these two fuels were studied and discussed, and the influencing factors like distillation temperature, kinematic viscosity and density were considered as important reasons. The result shows that when the surrogate fuel is injected into the ambient under supercritical conditions, the liquid penetration is slightly longer than that of No.2 diesel. The jet penetration of No.2 diesel is slightly longer than that of the surrogate fuel. Moreover, the spray cone angle of the surrogate fuel is wider than that of No.2 diesel. In summary, the fuels of different thermophysical characteristics may develop and diffuse differently resulting in different characteristics of injection and atomization.
Sun, Xiaochuan, Li, Xiang, Huang, Zhong, Ju, Dehao, Lu, Xing-cai, Han, Dong, Huang, Zhen
Application of Solid State Welding and Superplastic Forming to Aerospace Vehicles2017-01-21489/19/2017
In this paper, manufacturing of launcher structural components with superplastic forming (SPF) and solid state welding technologies is presented with several examples. Some of high strength aerospace alloys, like aluminum, titanium and superalloys, are known to have superplasticity so that complex shapes of aerospace components can be produced with this technology. A combination of superplastic forming and solid state welding processes produces lighter and stiffer components than one manufactured with conventional machining and welding. Solid state welding is an attractive method to weld materials without melting where mechanical properties are important since the welding interface is homogeneous without liquid phases. The results includes aluminum and titanium tanks cryogenic application, a scaled combustion chamber with duplex stainless steel and copper alloy, and lightweight aluminum-copper-lithium cylinder manufactured by superplastic forming and diffusion bonding or friction stir welding.
Lee, Ho-Sung
Multi-Pulse Motor (MPM) for Use with Electric Solid Propellants (ESP)TBMG-275389/1/2017
The purpose of this work was to create a safe, green, controllable solid rocket motor that can be pulsed a number of times with electricity to control the ignition and extinguishment times to produce a required thrust or impulse bit. The innovation features an Electric Solid Propellant (ESP). The key problem is that the geometry of the ESP grain changes because of the evolution of the propellant to exhaust gases, but a closed electrical circuit is required to keep the electrical power applied to the grain and continue burning. The chamber pressure is utilized in the multi-pulse motor (MPM) design to ensure electrical contact is retained during the pulsing event.
Combustion Characteristics of Hydrocarbon Droplets Induced by Photoignition of Aluminum NanoparticlesTBMG-273818/1/2017
In the study of combustion characteristics of liquid rocket fuels, it is customary to either study the combustion of liquid fuel droplets or the combustion of fuel sprays. However, the two are closely related to each other, because in a typical rocket combustion chamber, the burning of droplets, droplet clusters, and fuel sprays occur simultaneously.
Measuring Propellant Stress Relaxation Modulus Using Dynamic Mechanical AnalyzerTBMG-273808/1/2017
Structural analysis of solid rocket motors is challenging for several reasons, but the most important of these is the complex behavior of the propellant. The mechanical response of a solid propellant is time and temperature dependent. The complexity of the mathematical analysis of the propellant depends on the loading conditions, but for some loading situations, the linear viscoelasticity assumption is reasonable. In particular, linear viscoelasticity is perhaps the most appropriate material behavior description for use in the simulations of stresses related to storage conditions. Typically, simulations use a viscoelastic model in the form of a Prony series and a Williams–Landel–Ferry (WLF) equation. The parameters in these models are derived from stress relaxation experiments, making the stress relaxation experiment a key viscoelastic test, analogous to the tensile test for linear elastic materials.
Measuring Propellant Stress Relaxation Modulus Using Dynamic Mechanical Analyzer17AERP08_066/1/2017
New testing technique requires less material, gives more accurate results. Air Force Research Laboratory (AFMC), Edwards Air Force Base, California Structural analysis of solid rocket motors is challenging for several reasons, but the most important of these is the complex behavior of the propellant. The mechanical response of a solid propellant is time and temperature dependent. The complexity of the mathematical analysis of the propellant depends on the loading conditions, but for some loading situations, the linear viscoelasticity assumption is reasonable. In particular, linear viscoelasticity is perhaps the most appropriate material behavior description for use in the simulations of stresses related to storage conditions. Typically, simulations use a viscoelastic model in the form of a Prony series and a Williams-Landel-Ferry (WLF) equation. The parameters in these models are derived from stress relaxation experiments, making the stress relaxation experiment a key viscoelastic test, analogous to the tensile test for linear elastic materials. A typical set of stress relaxation tests is performed at several discrete temperatures that cover a range of temperatures anticipated by the fielded motor. At each of the selected temperatures, the specimen is deformed with approximately a single step in strain, which is then held constant for the duration of the test. While held at this constant strain, the stress decays over time due to relaxation of the rubbery elastomer. During this portion of the test, the stresses are measured, and the ratio of stress to applied strain is determined. This ratio is termed the stress relaxation modulus ER. Using time-temperature superposition, the set of curves at the various temperatures can be shifted horizontally relative to each other to form a master curve. The translation of the curves takes a specific mathematical form, viz., the WLF equation. From this master curve, the Prony series at any given temperature can be calculated, and the calculation can be incorporated into finite element analyses along with the WLF equation, making linear viscoelastic analysis of rocket motors possible.
Combustion Characteristics of Hydrocarbon Droplets Induced by Photoignition of Aluminum Nanoparticles17AERP08_076/1/2017
Test methodology allows analysis of combustion dynamics for subscale rocket injectors under super critical conditions. Air Force Research Laboratory (AFMC), Edwards Air Force Base, California In the study of combustion characteristics of liquid rocket fuels, it is customary to either study the combustion of liquid fuel droplets or the combustion of fuel sprays. However, the two are closely related to each other, because in a typical rocket combustion chamber, the burning of droplets, droplet clusters, and fuel sprays occur simultaneously. The study of droplet burning is mostly focused either on single suspended droplets or free droplets, usually in a free fall. In either case, the conditions regarding the combustion of a single fuel droplet deviates from what is expected in typical liquid rocket engines (LREs). This not only concerns the chamber pressure and the temperature that is usually much lower in a typical droplet study, but also the droplet ignition process and the related transient effects such as temperature ramp up that is much slower than what takes place in LREs. As a result, reported values of combustion parameters such as the burning rate constant and the ignition delay time obtained under typical experimental conditions may not entirely represent the combustion processes in LREs.
Green Monopropellant Secondary Payload Propulsion SystemTBMG-268355/1/2017
Small satellites, launched as secondary payloads, are increasingly being fielded. Advances in liquid rocket propulsion that enhance the on-orbit maneuverability, increase the on-orbit life, and decrease the time between identified need for and deployment of such spacecraft are of great value. Replacing the nearly ubiquitous yet toxic hydrazine propellant with AF-M315E produces higher specific impulse and density specific impulse, resulting in improved overall velocity change capability and increased on-orbit life.
Low Solidity Vaned Diffuser (LSVD) Design for Improvement of Pressure RecoveryTBMG-268525/1/2017
Many pump vaned diffuser designs are based on existing airfoil designs, with little attention given to the vane leading edge. There is a need for a vaned diffuser leading edge that helps resist flow separation and the resultant poor diffuser pressure recovery. Diffusers in pumps are often working with an incompressible fluid that makes potential flow methodologies — which have incompressibility as a boundary condition — attractive. The potential flow-based free-streamline analysis methods have been known to improve the aerodynamics of varied components at high incidence angles, such as diffusers, jet engine nacelles, and liquid rocket engine turbopump inducers.
Electric Rockets and the Future of Satellite Propulsion17AERP04_034/1/2017
Humans have been using rocket propulsion for almost a millennium, starting with Chinese rockets and “fire arrows” in the 13th century and continuing to the modern era's powerful Space Shuttle and Falcon rockets. For most of that history, rockets have been chemically fueled, but in the past century scientists and engineers have also experimented with electric rockets, also known as ion engines or ion propulsion systems. Rather than using chemical reactions to create heat and accelerate a propellant, electric rockets use electromagnetic or electrostatic fields acting on charged ions of propellant, speeding them up and shooting them out, away from the vehicle, producing thrust. The electrical energy to generate these fields comes from the sun, from batteries, or both.
Cooling Solution Helps NASA Get Closer to MarsTBMG-266484/1/2017
Cold conditioning system Aggreko Houston, TX 281-985-8200
Electric Rockets and the Future of Satellite PropulsionTBMG-266794/1/2017
Humans have been using rocket propulsion for almost a millennium, starting with Chinese rockets and “fire arrows” in the 13th century and continuing to the modern era's powerful Space Shuttle and Falcon rockets. For most of that history, rockets have been chemically fueled, but in the past century scientists and engineers have also experimented with electric rockets, also known as ion engines or ion propulsion systems.
A Modular Apparatus and Method for Attaching Multiple DevicesTBMG-266464/1/2017
Scientists at NASA's Glenn Research Center have received several patents for fabrication methods that speed the production of silicon carbide (SiC) sensors and electronics, reduce fabrication costs, and enable advanced semiconductor functionality at temperatures greater than 600 °C. Glenn has developed a method for fabricating ultra-thin SiC microstructures using a dopant selective reactive ion etching (DSRIE) technique that can create extremely thin diaphragms (approximately 2 microns), increasing sensor sensitivity and resolution. In addition, Glenn has developed a modular protective packaging that allows SiC-based electronic devices to survive and operate reliably at very high temperatures. These innovations improve the real-time monitoring of high-temperature harsh environments, such as jet and rocket engines, allowing faster response times and more accurate readings.
Experimental Measurements and Computations for Clarifying Nearly Complete Air-Insulation Obtained by the Concept of Colliding Pulsed Supermulti-Jets2017-01-10303/28/2017
In our previous papers, a new concept of a compressive combustion engine (Fugine) was proposed based on the collision of pulsed supermulti-jets, which can enclose the burned gas around the chamber center leading to an air-insulation effect and also a lower exhaust gas temperature due to high single-point compression. In order to examine the compression level and air-insulation effect as basic data for application to automobiles, aircraft, and rockets, a prototype engine based on the concept, i.e., a piston-less prototype engine with collision of bi-octagonal pulsed multi-jets from fourteen nozzles, was developed. Some combustion results [Naitoh et al. SAE paper, 2016] were recently reported. However, there was only one measurement of wall temperature and pressure in the previous report. Thus, in this paper, more experimental data for pressures and temperatures on chamber walls and exhaust temperatures, are presented for the prototype engine. First, pressure over 0.6MPa was measured on the chamber wall. A nearly complete air insulation effect was presumably obtained based on the experimental data for temperature measured on the chamber wall. The measured exhaust temperature was at an intermediate level around 700K. Experimental data are also presented for the air-insulation effect on a small solid wall located downstream from the collision point of the supermulti-jets. Unsteady three-dimensional computations of compressible flow also indicate that the experimental result of 0.6 MPa at the cylinder wall implies pressure of about 5 MPa at the collision point of the jets. The potential for high thermal efficiency is evaluated on the basis of the data.
Konagaya, Remi, Oyanagi, Susumu, Kanase, Takuto, Tsuchiya, Jumpei, Ayukawa, Ken, Kinoshita, Kodai, Mikoda, Junya, Fujita, Hirotaka, Naitoh, Ken
Study of Formulating Conserved Scalar Equations for Turbulent Reactive Flows with General Species Mass- Diffusion Coefficients for Utilization in Flamelet ModelsTBMG-261651/1/2017
The modeling of turbulent reactive flows is a subject of contemporary research. Current turbulent-reaction models cannot account for realistic complexities such as distinct species mass-diffusion coefficients. Under the assumption of a single, constant, mass-diffusion coefficient, a conserved-scalar equation is typically derived in turbulent reactive flows by taking the difference between chemical-species conservation equations having opposite reaction rates (in the sense that the reactant has an opposite reaction rate to the product), thereby creating an equation devoid of reaction terms. Assuming the reaction regions are very thin and are merely contorted by turbulence, chemistry and turbulence can be decoupled, and the evolving statistics of the conserved scalar describe the reaction progression. No such equation has yet been derived for distinct mass-diffusion coefficient cases where the single coefficient is now replaced by a full matrix. Considering that mass diffusion is responsible for reactants approaching at the molecular level and for reaction initiation, this lack of mathematical framework is very disturbing.
NASA's Software Answers Big Questions About Space and EarthTBMG-2606312/1/2016
NASA’s next-generation Space Launch System (SLS) for deep space exploration consists of four RS-25 liquid rocket engines and two five-segment solid rocket boosters (SRBs). At ignition, the SRBs create a significant over-pressure event, known as the Ignition Over-Pressure (IOP) event. The IOP event experienced in the first space shuttle flight (STS-1) in 1981 damaged the space shuttle orbiter and motivated significant design changes to the launch pad at NASA’s Kennedy Space Center. The RS-25 liquid rocket engines aboard the SLS, along with the SRBs, also create a very significant acoustic environment that, if not mitigated, can harm the SLS launch vehicle. An Ignition Over-Pressure/Sound Suppression (IOP/SS) water system was developed to reduce the amplitude of both the IOP event and the acoustic environment caused by the firing of the RS-25 engines and the SRBs.
Uncertainty of In-Flight Thrust DeterminationAIR1678B (Current)10/22/2016
This document defines and illustrates the process for determination of uncertainty of turbofan and turbojet engine in-flight thrust and other measured in-flight performance parameters. The reasons for requiring this information, as specified in the E-33 Charter, are: determination of high confidence aircraft drag; problem rectification if performance is low; interpolation of measured thrust and aircraft drag over a range of flight conditions by validation and development of high confidence analytical methods; establishment of a baseline for future engine modifications. This document describes systematic and random measurement uncertainties and methods for propagating the uncertainties to the more complicated parameter, in-flight thrust. Methods for combining the uncertainties to obtain given confidence levels are also addressed. Although the primary focus of the document is in-flight thrust, the statistical methods described are applicable to any measurement process. The E-33 Committee has endeavoured to gather industry-wide expertise in in-flight measurement and uncertainty analysis to collect and promulgate recommended practices in the subject disciplines. The Committee is organized into subcommittees to address both the analytical and test methodology for determination of in-flight thrust and also the uncertainty of the determination. This document; Uncertainty of In-flight Thrust Determination, AIR1678, addresses the process for determining the uncertainty of in-flight thrust. A companion document, In-Flight Thrust Determination, AIR1703, addresses the basic methodology for determining in-flight thrust. The Committee, after reviewing recommended changes and clarification in definitions and application of statistical uncertainty items, made small revisions to the original document published in 1985. These changes were incorporated into AIR1678 Rev A. This Revision B has the same Scope as preceding versions. The nomenclature and methodology used herein are now consistent with evolving world and national standards promulgated primarily by ISO and ASME.
E-33 In Flight Propulsion Measurement Committee
Magnetostrictive Pressure Regulating SystemTBMG-2551410/1/2016
NASA’s Marshall Space Flight Center has developed a set of unique magnetostrictive (MS) technologies for utilization in pressure regulation and valve systems. By combining MS-based sensors with a newly designed MS-based valve, Marshall has developed an advanced MS regulator. This innovative approach provides both a regulator and a valve with rapid response times. In addition, the components are lightweight, compact, highly precise, and can operate over a wide range of temperatures and pressures. A prototype of the MS valve has been developed and NASA is seeking partners for licensure of this novel technology.
Seeing Closer and Clearer with NASA Imaging TechnologyTBMG-2553810/1/2016
NASA’s Origins, Spectral Interpretation, Resource Identification, Security-Regolith Explorer (OSIRIS-Rex) spacecraft launched on September 8 to the near-Earth asteroid Bennu to harvest a sample of surface material and return it to Earth for study. But before the science team selects a sample site, they can find out a bit about Bennu’s elemental make-up.
Eddy-Current-Minimizing Flow Plug for Use in Flow Conditioning and Flow MeteringTBMG-2551010/1/2016
Innovators at NASA’s Marshall Space Flight Center have developed a suite of prototype fluid plug technologies with an array of capabilities for fluid flow metering, mixing, and conditioning. Each innovation within this suite is based upon a core technology that has no moving parts, is simple to manufacture, and provides high reliability and efficiency. Also, the base fluid plug technology can be modified with very few or no hardware changes to achieve the desired effect or combination of mixing, metering, and conditioning capabilities depending on the application.
Actuation System Data Summary for Missiles and Launch VehiclesAIR1855A (Current)9/28/2016
This SAE Aerospace Information Report (AIR) includes all missile and launch vehicle actuation systems, including electrohydraulic, electropneumatic, and electromechanical types. The data for many systems are not complete. As more information becomes available, periodic updates will be issued to complete existing data sheets and to add new ones. An index by type of vehicle and by type of actuation system is included. The actual data sheets in the body of the report are organized in alphabetical order.
A-6B1 Hydraulic Servo Actuation Committee
Ultralight Self-Deployable Solar SailsTBMG-253469/1/2016
Deployment of large structures such as solar sails relies typically upon electromechanical mechanisms, mechanically expandable or inflatable booms, launch restraints, controls, and other mechanisms that drastically increase the total mass, stowage volume, and areal density. The primary performance parameter for solar sails is areal density, which determines the acceleration of the sail. Present technology allows the solar sail areal density to be around 20 g/m2, and that permits only nearby demonstration missions.
Hermetic Seal Leak Detection Apparatus with Variable Size Test ChamberTBMG-249287/1/2016
NASA’s Marshall Space Flight Center has developed a unique apparatus ideal for use in nondestructive testing (NDT) of hermetic seals of containers or instrumentation. The device is capable of detecting both large and small leaks and can be calibrated to characterize the relative leak rate. Its simple design does not require specialized gases for pressurization and detection, and eliminates the need for expensive instrumentation such as a mass spectrometer to analyze leaks and achieve high sensitivity. Low in cost and simple to manufacture, the patent-pending technology is ideal for use in many industries, from aerospace applications to food packaging and commercial goods.
Method for Manufacturing a Thin Film Structural SystemTBMG-249627/1/2016
NASA’s Langley Research Center has developed a technology that uses commercially available additive print manufacturing to add various levels of structural hierarchy to thin-film surfaces. The approach adds very little mass to thin films, but provides substantial performance enhancements, such as increased damage tolerance to tearing and ripping. NASA developed this technology to provide new and improved ways to produce robust, ultra-lightweight space structures such as solar sails, solar shades, and antennas. Beyond space applications, the technology is well suited for other thin-film applications.
Rangefinder for Measuring Volume of Cryogenic or Caustic Turbulent FluidsTBMG-248376/1/2016
Specific impulse (ISP), or simply impulse (change in momentum) per unit amount of propellant consumed, is a measure of rocket and jet engine efficiency. The amount of propellant, or in the case of engine testing at the Stennis Space Center (SSC), cryogen consumed during rocket engine testing must be measured to accurately quantify ISP. One way to determine the amount of cryogen used is to measure the change in cryogen fluid height within a storage/feed tank during testing and then relate the change in height to volume of cryogen consumed. A float system coupled with discrete vertically positioned Reed switches is currently used at the SSC to determine cryogen fluid height and then determine cryogen consumed during a rocket motor test firing. However, the cryogen fluid level within a run tank varies continuously and the switches are placed at discrete locations, limiting the accuracy of this method. If individual switch failures occur, the error increases due to the increased distance between switches/measurement locations. In addition, since pressurized gas is used to force the significantly cooler liquid cryogen out of the tank during a test, the liquid cryogen surface is turbulent and not flat or smooth, which can also affect accuracy.
Advanced Magnetostrictive Regulator, Valve, and Force-to-Angle SensorTBMG-245305/1/2016
Typical aerospace rocket engines use valves to control the flow and pressures of propellant and pressurants. These typical valves are designed to operate with a mechanical, electromechanical, or pneumatic operator. They all have at least one, and often multiple, penetrations from the fluid to the operator’s prime mover. The penetrations are sources for leaks, failures, and are often considered to be unreliable for use in single string systems. Therefore, the fluid system designer frequently will utilize several parallel path valves, effectively doubling the resources needed to accomplish the task. These redundant valves allow for isolation of the potentially leaking fluid penetrations. If the systems cannot afford the multiple path approach, then the valves are subjected to high levels of testing and quality control, or utilize bellows or other expensive and difficult to handle/design and costly features.
The Future of Exploration Starts With 3D PrintingTBMG-246595/1/2016
Last year, engineers at NASA’s Marshall Space Flight Center in Huntsville, AL, tested an additive manufacturing process that is being used to make some of the parts for NASA’s new rocket, the Space Launch System (SLS), more efficiently and affordably without compromising performance and safety. Selective laser melting (SLM) is a 3D printing process used to create complex parts for the engine and other components of the rocket. Four RS-25 engines and two solid rocket boosters will power the core stage of the SLS.
Distributed, Fast, Intelligent Infrastructure Health Monitoring SystemTBMG-242674/1/2016
NASA needed an innovative solution to conduct system assessments of frameworks and to characterize subsystems of interest. This tool would be required to determine anomalies; examine their causes (root-cause analysis); make predictive statements (prognostics); provide intelligent health monitoring with incremental knowledge for working with unknown scenarios; and provide maintenance support for a complex collection of systems, subsystems, and elements in rocket engine test platforms.
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