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At least 109 records · Page 6

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NASA NTRS1974

Rapid estimation of characteristics of gas dynamic lasers

Sudden-freeze approximation is applied to the flow of a CO2-N2-He mixture in wedge-type nozzles. This approximation permits rapid estimation of the freezing temperature of the upper laser level as a function of the stagnation pressure and the nozzle geometry. The stagnation temperature and the composition of the mixture appear as parameters. Gain and power output may then be estimated and calculations are presented for two cases.

Murty, S. S. R.↗

NASA NTRS1996

Mach 5 to 7 RBCC Propulsion System Testing at NASA-LeRC HTF

A series of Mach 5 to 7 freejet tests of a Rocket Based Combined Cycle (RBCC) engine were cnducted at the NASA Lewis Research Center (LERC) Hypersonic Tunnel Facility (HTF). This paper describes the configuration and operation of the HTF and the RBCC engine during these tests. A number of facility support systems are described which were added or modified to enhance the HTF test capability for conducting this experiment. The unfueled aerodynamic perfor- mance of the RBCC engine flowpath is also presented and compared to sub-scale test results previously obtained in the NASA LERC I x I Supersonic Wind Tunnel (SWT) and to Computational Fluid Dynamic (CFD) analysis results. This test program demonstrated a successful configuration of the HTF for facility starting and operation with a generic RBCC type engine and an increased range of facility operating conditions. The ability of sub-scale testing and CFD analysis to predict flowpath performance was also shown. The HTF is a freejet, blowdown propulsion test facility that can simulate up to Mach 7 flight conditions with true air composition. Mach 5, 6, and 7 facility nozzles are available, each with an exit diameter of 42 in. This combination of clean air, large scale, and Mach 7 capabilities is unique to the HTF. This RBCC engine study is the first engine test program conducted at the HTF since 1974.

Perkins, H. Douglas↗

NASA NTRS2009

Advances in Non-contact Measurement of Creep Properties

Our team has developed a novel approach to measuring creep at extremely high temperatures using electrostatic levitation (ESL). This method has been demonstrated on niobium up to 2300 C, while ESL has melted tungsten (3400 C). High-precision machined spheres of the sample are levitated in the NASA MSFC ESL, a national user facility, and heated with a laser. The laser is aligned off-center so that the absorbed photons transfer their momentum to the sample, causing it to rotate at up to 250,000+ RPM. The rapid rotation loads the sample through centripetal acceleration, causing it to deform. The deformation of the sample is captured on high-speed video, which is analyzed by machine-vision software from the University of Massachusetts. The deformations are compared to finite element models to determine the constitutive constants in the creep relation. Furthermore, the noncontact method exploits stress gradients within the sample to determine the stress exponent in a single test. This method was validated in collaboration with the University of Tennessee for niobium at 1985 C, with agreement within the uncertainty of the conventional measurements. A similar method is being employed on Ultra-High-Temperature ZrB2- SiC composites, which may see application in rocket nozzles and sharp leading edges for hypersonic vehicles.

Hyers, Robert W.↗

NASA NTRS2009

Liquid Engine Design: Effect of Chamber Dimensions on Specific Impulse

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.

Hoggard, Lindsay↗

NASA NTRS1993

Micromechanical simulation of damage progression in carbon phenolic composites

Carbon/phenolic composites are used extensively as ablative insulating materials in the nozzle region of solid rocket motors. The current solid rocket motor (RSRM) on the space shuttle is fabricated from woven rayon cloth which is carbonized and then impregnated with the phenolic resin. These plies are layed up in the desired configuration and cured to form the finished part. During firing, the surface of the carbon/phenolic insulation is exposed to 5000 F gases from the rocket exhaust. The resin pyrolizes and the material chars to a depth which progresses with time. The rate of charring and erosion are generally predictable, and the insulation depth is designed to allow adequate safety margins over the firing time of the motor. However, anomalies in the properties and response of the carbon/phenolic materials can lead to severe material damage which may decrease safety margins to unacceptable levels. Three macro damage modes which were observed in fired nozzles are: ply lift, 'wedge out', and pocketing erosion. Ply lift occurs in materials with plies oriented nearly parallel to the surface. The damage occurs in a region below the charred material where material temperatures are relatively low - about 500 F. Wedge out occurs at the intersection of nozzle components whose plies are oriented at about 45 deg. The corner of the block of material breaks off along a ply interface. Pocketing erosion occurs in material with plies oriented normal to the surface. Thermal expansion is restrained in two directions resulting in large tensile strains and material failure normal to the surface. When a large section of material is removed as a result of damage, the insulation thickness is reduced which may lead to failure of the nozzle due to excessive heating of critical components. If these damage events cannot be prevented with certainty, the designer must increase the thickness of the insulator thus adding to both weight and cost. One of the difficulties in developing a full understanding of these macro damage mechanisms is that the loading environment and the material response to that environment are extremely complex. These types of damage are usually only observed in actual motor firings. Therefore, it is difficult and expensive to evaluate the reliability of new materials. Standard material tests which measure mechanical and thermal properties of test specimens can only provide a partial picture of how the material will respond in the service environment. The development of the ANALOG test procedure which can combine high heating rates and mechanical loads on a specimen will improve the understanding of the interactive effects of the various loads on the system. But a mechanistic model of material response which can account for the heterogeneity of the material, the progression of various micromechanical damage mechanisms, and the interaction of mechanical and thermal stresses on the material is required to accurately correlate material tests with response to service environments. A model based on fundamental damage mechanisms which is calibrated and verified under a variety of loading conditions will provide a general tool for predicting the response of rocket nozzles. The development of a micromechanical simulation technique was initiated and demonstrated to be effective for studying across-ply tensile failure of carbon/phenolic composites.

Slattery, Kerry T.↗

NASA NTRS1995

Mass spectrometer measurements of test gas composition in a shock tunnel

Shock tunnels afford a means of generating hypersonic flow at high stagnation enthalpies, but they have the disadvantage that thermochemical effects make the composition of the test flow different to that of ambient air. The composition can be predicted by numerical calculations of the nozzle flow expansion, using simplified thermochemical models and, in the absence of experimental measurements, it has been necessary to accept the results given by these calculations. This note reports measurements of test gas composition, at stagnation enthalpies up to 12.5 MJ.kg(exp -1), taken with a time-of-flight mass spectrometer. Limited results have been obtained in previous measurements. These were taken at higher stagnation enthalpies, and used a quadruple mass spectrometer. The time-of-flight method was preferred here because it enabled a number of complete mass spectra to be obtained in each test, and because it gives good mass resolution over the range of interest with air (up to 50 a.m.a.).

Skinner, K. A.↗

NASA NTRS2012

Real-Time X-ray Radiography Diagnostics of Components in Solid Rocket Motors

Solid rocket motors (SRMs) typically use nozzle materials which are required to maintain their shape as well as insulate the underlying support structure during the motor operation. In addition, SRMs need internal insulation materials to protect the motor case from the harsh environment resulting from the combustion of solid propellant. In the nozzle, typical materials consist of high density graphite, carbon-carbon composites and carbon phenolic composites. Internal insulation of the motor cases is typically a composite material with carbon, asbestos, Kevlar, or silica fibers in an ablative matrix such as EPDM or NBR. For both nozzle and internal insulation materials, the charring process occurs when the hot combustion products heat the material intensely. The pyrolysis of the matrix material takes away a portion of the thermal energy near the wall surface and leaves behind a char layer. The fiber reinforcement retains the porous char layer which provides continued thermal protection from the hot combustion products. It is of great interest to characterize both the total erosion rates of the material and the char layer thickness. By better understanding of the erosion process for a particular ablative material in a specific flow environment, the required insulation material thickness can be properly selected. The recession rates of internal insulation and nozzle materials of SRMs are typically determined by testing in some sort of simulated environment; either arc-jet testing, flame torch testing, or subscale SRMs of different size. Material recession rates are deduced by comparison of pre- and post-test measurements and then averaging over the duration of the test. However, these averaging techniques cannot be used to determine the instantaneous recession rates of the material. Knowledge of the variation in recession rates in response to the instantaneous flow conditions during the motor operation is of great importance. For example, in many SRM configurations the recession of the solid propellant grain can drastically alter the flow-field and effect the recession of internal insulation and nozzle materials. Simultaneous measurement of the overall erosion rate, the development of the char layer, and the recession of the char-virgin interface during the motor operation can be rather difficult. While invasive techniques have been used with limited success, they have serious drawbacks. Break wires or make wire sensors can be installed into a sufficient number of locations in the charring material from which a time history of the charring surface can be deduced. These sensors fundamentally alter the local structure of the material in which they are imbedded. Also, the location of these sensors within the material is not known precisely without the use of an X-ray. To determine instantaneous recession rates, real-time X-ray radiography (X-ray RTR) has been utilized in several SRM experiments at PSU. The X-ray RTR system discussed in this paper consists of an X-ray source, X-ray image intensifier, and CCD camera connected to a capture computer. The system has been used to examine the ablation process of internal insulation as well as nozzle material erosion in a subscale SRM. The X-ray source is rated to 320 kV at 10 mA and has both a large (5.5 mm) and small (3.0 mm) focal spot. The lead-lined cesium iodide X-ray image intensifier produces an image which is captured by a CCD camera with a 1,000 x 1,000 pixel resolution. To produce accurate imagery of the object of interest, the alignment of the X-ray source to the X-ray image intensifier is crucial. The image sequences captured during the operation of an SRM are then processed to enhance the quality of the images. This procedure allows for computer software to extract data on the total erosion rate and the char layer thickness. Figure 1 Error! Reference source not found.shows a sequence of images captured during the operation the subscale SRM with the X-ray RTR system. The X-rayTR system, alignment procedure, uncertainty determination, and image analysis process will be discussed in detail in the full manuscript.

Cortopassi, A. C.↗

NASA NTRS1990

Multi-element spherical shell generation

A nozzle assembly in a multi-element spherical shell generation system includes first and second side-by-side spaced apart nozzles and a web portion extending between and connecting the nozzles. The first nozzle has an inner orifice adapted to discharge a first filler material and an outer annular orifice separated from and defined in concentric relation about the inner orifice and adapted to discharge a first shell material. The second nozzle has an inner orifice adapted to discharge a second filler material and an outer annular orifice separated from and defined in concentric relation about the inner orifice and adapted to discharge a second shell material. A multi-element spherical shell can be formed through employment of the nozzle assembly by merger with one another after discharge from the outer orifices of the nozzles of a pair of adjacent annular streams of liquid or molten shell wall material of different compositions and encapsulation by the mixed shell wall materials of a common encapsulated core fluids also simultaneously discharged by the inner orifices nozzles. On the other hand, the pair of encapsulating streams of shell wall material can be of the same materials which merge together and encapsulate core fluids of different compositions which will merge together after discharge from the nozzles.

Morrison, Andrew D.↗

NASA NTRS2017

Creep and Cyclic Fatigue Durability of 3D Woven SiC/SiC Composites with (CVI+PIP) Hybrid Matrix

SiC-SiC composites are potential candidate materials for turbine components such as combustor liners, nozzle vanes and blades because of their low density, high temperature capability, and tailorable mechanical properties. The first generation of SiC-SiC composites fabricated by melt infiltration are being introduced in current engines. These materials are limited to 1315 degrees Centigrade applications because of the presence of residual silicon in the SiC matrix. Currently there is an increasing interest in developing SiC-SiC composites without silicon for structural aerospace applications above 1315 degrees Centigrade. The developmental 3D woven SiC-SiC composites with Chemical Vapor Infiltration plus Polymer Impregnation and Pyrolysis (CVI+PIP) hybrid matrix show potential for 1482 degrees Centigrade applications. In this study 3D woven SiC-SiC composite specimens were creep tested at 1482 degrees Centigrade at different stress levels until failure or for 1000 hours without failure. The failed specimens were analyzed for under a scanning electron microscope to assess the damage mechanisms. The creep data of 3D woven SiC-SiC composites with (CVI+PIP) hybrid matrix were compared with those of full CVI SiC-SiC composites and sintered SiC from the literature. The potential benefits limitations and durability of 3D woven SiC-SiC composites for turbine applications will be discussed.

creep durability↗

NASA NTRS1969

Natural gas flow through critical nozzles

Empirical method for calculating both the mass flow rate and upstream volume flow rate through critical flow nozzles is determined. Method requires knowledge of the composition of natural gas, and of the upstream pressure and temperature.

Johnson, R. C.↗

NASA NTRS1980

Technology status of a liquid fluorine-hydrazine rocket engine for a planetary spacecraft propulsion system

This paper discusses the current status of a fluorine-hydrazine rocket engine development program. Incorporation of a thin rhenium inner liner successfully eliminated corrosion of the carbon/carbon composite thrust chamber wall experienced during a previous test program. The results of hot-fire tests utilizing reworked and new injectors which provide increased fuel film cooling showed that thrust chamber head-end temperatures could be maintained at an acceptable level. As expected, the accompanying specific impulse performance loss requires optimizing the amount of film cooling to minimize the loss. The efforts to refine the rhenium liner vapor deposition process culminated in a carbon/carbon composite thrust chamber total test duration of 1008 seconds. Tasks presently in process include: (1) fabrication of two carbon/carbon composite thrust chambers incorporating 60:1 expansion ratio nozzles; (2) injector tests to optimize performance and cooling; (3) additional refinements to the rhenium lining process; and (4) fabrication and test of a freestanding rhenium thrust chamber.

Appel, M. A.↗

NASA NTRS1991

Propulsion system needs

The needs of the designer of a solid core nuclear rocket engine are discussed. Some of the topics covered include: (1) a flight thrust module/feed system module assembly; (2) a nuclear thermal rocket (NTR), expander cycle, dual T/P; (3) turbopump operating conditions; (4) typical system parameters; (5) growth capability composite fuel elements; (6) a NTR radiation cooled nozzle extension; (7) a NFS-3B Feed System; and (8) a NTR Integrated Pneumatic-Fluidics Control System.

Gunn, Stanley↗

NASA NTRSDate not supplied

Lightweight Thrust Chamber Assemblies using Multi-Alloy Additive Manufacturing and Composite Overwrap

Additive Manufacturing (AM) has brought significant design and fabrication opportunities for complex components with internal features such as liquid rocket engine thrust chambers not previously possible. This technology allows for significant cost savings and schedule reductions in addition to new performance optimization through weight reduction and increased margins. Specific to regeneratively-cooled combustion chambers and nozzles for liquid rocket engines, additive manufacturing offers the ability to form the complex internal coolant channels and the closeout of the channels to contain the high pressure liquid propellants with a single operation. Much of additive manufacturing development has focused on monolithic alloys using Laser Powder Bed Fusion (L-PBF), which do not allow for complete optimization of the structure. The National Aeronautics and Space Administration (NASA) completed feasibility of an AM bimetallic L-PBF GRCop-84 copper-alloy combustion chamber with an AM electron beam freeform Inconel 625 structural jacket under the Low Cost Upper Stage Propulsion (LCUSP) Project. A follow-on project called Rapid Analysis and Manufacturing Propulsion Technology (RAMPT) is under development to further expand large-scale multi-alloy thrust chambers while maturing composite overwrap technology for significant weight savings opportunities. The RAMPT project has three primary objectives: 1) Advancing blown powder Directed Energy Deposition (DED) to fabricate integral-channel large scale nozzles, 2) Develop composite overwrap technology to reduce weight and provide structural capability for thrust chamber assemblies, and 3) Develop bimetallic and multi-metallic additively manufactured radial and axial joints to optimize material performance. In addition to these primary manufacturing developments, analytical modeling efforts compliment the process development to simulate the AM processes to reduce build failures and distortions. The RAMPT project is also maturing the supply chain for various manufacturing processes described above in addition to L-PBF of GRCop-42. This paper will present an overview of the RAMPT project, the process development and hardware progress to date, material and hot-fire testing results, along with future developments.

Additive Manufacturing↗

NASA NTRS1974

Composition surveys of test gas produced by a hydrogen-oxygen-air burner

As a result of the need for a uniform hot gas test stream for fuel injector development for hydrogen fueled supersonic combustion ramjet engines, an experimental study of injector configuration effect on exit flow uniformity of a hydrogen fueled oxygen replenished, combustion burner was made. Measurements used to investigate the burner nozzle exit profiles were pitot and gas sample measurements. Gas composition and associated temperature profiles were reduced to an acceptable level by burner injector modifications. The effect of the injector modifications was to redistribute the hydrogen fuel, increase the air pressure drop, promote premixing of the oxygen and air, and establish a uniform flow pattern where the oxygen-air mixture comes into contact with the hydrogen fuel. The most sensitive phenomenon which affected the composition profiles was the uniformity of the air distribution supplied to the combustion chamber.

Eggers, J. M.↗

NASA NTRS2016

Influence of Constituents on Creep Properties of SiC/SiC Composites

SiC-SiC composites are being considered as potential candidate materials for next generation turbine components such as combustor liners, nozzle vanes and blades because of their low density, high temperature capability, and tailorable mechanical properties. These composites are essentially fabricated by infiltrating matrix into a stacked array of fibers or fiber preform by one or a combination of manufacturing methods such as, Melt Infiltration (MI) of molten silicon metal, Chemical Vapor Infiltration (CVI), Polymer Infiltration and Pyrolysis (PIP). To understand the influence of constituents, the SiC-SiC composites fabricated by MI, CVI, and PIP methods were creep tested in air between 12000 and 14500 degrees Centigrade for up to 500 hours. The failed specimens were analyzed under a scanning electron microscope to assess damage mechanisms. Also, knowing the creep deformation parameters of the fiber and the matrix under the testing conditions, the creep behavior of the composites was modeled and compared with the measured data. The implications of the results on the long term durability of these composites will be discussed.

modeling↗

NASA NTRS1981

Space Shuttle SRM performance improvement

A number of options for Solid Rocket Motor (SRM) performance improvement were studied and evaluated by NASA with respect to payload improvement potential, technical risk, schedule, cost, and impact on Shuttle system performance. It was found that the SRM delivered specific impulse could be improved by increasing the nozzle expansion ratio from 7.16 to 7.72. This was achieved by decreasing the throat diameter of the nozzle exit cone, while staying within tooling, facility, and ICD limitations. A simple modification of the grain inhibiting pattern was made to reshape the thrust-time history and thereby obtain more delivered impulse during the early portion of motor operation. These performance options, initiated in October 1980, provide a 3,000 lb payload increase. Longer range options are the use of a filament wound composite case, a HTPB propellant, and further increases in the nozzle expansion ratio. These improvements have a potential of 10,000 to 11,000 lb payload increase.

Speak, C. A.↗

NASA NTRS1981

Navier-Stokes solutions for nonaxisymmetric nozzle flows

The compressible Navier-Stokes equations for two-dimensional and three-dimensional flow are solved for nonaxisymmetric nozzle configurations. Both internal flow alone and combined internal and external flow cases are considered. An implicit approximate factorization algorithm is employed to solve the governing equations which are cast in terms of a body-fitted coordinate system. Turbulence closure is accomplished with a composite algebraic model. Two-dimensional results for a wedge plug nozzle flow in the form of surface pressures compare favorably with experimental data. Three-dimensional calculations for two nonaxisymmetric nozzles are also presented and compared with data.

Swanson, R. C.↗

NASA NTRS1982

Energy efficient engine high pressure turbine test hardware detailed design report

The high pressure turbine configuration for the Energy Efficient Engine is built around a two-stage design system. Moderate aerodynamic loading for both stages is used to achieve the high level of turbine efficiency. Flowpath components are designed for 18,000 hours of life, while the static and rotating structures are designed for 36,000 hours of engine operation. Both stages of turbine blades and vanes are air-cooled incorporating advanced state of the art in cooling technology. Direct solidification (DS) alloys are used for blades and one stage of vanes, and an oxide dispersion system (ODS) alloy is used for the Stage 1 nozzle airfoils. Ceramic shrouds are used as the material composition for the Stage 1 shroud. An active clearance control (ACC) system is used to control the blade tip to shroud clearances for both stages. Fan air is used to impinge on the shroud casing support rings, thereby controlling the growth rate of the shroud. This procedure allows close clearance control while minimizing blade tip to shroud rubs.

Halila, E. E.↗