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NASA NTRS1974

Space shuttle maneuvering engine reusable thrust chamber program. Task 11: Low Epsilon stability test plan

The performance and heat transfer characteristics of a doublet element type injector for the space shuttle orbiter maneuvering engine thrust chamber were investigated. Ths stability characteristics were evaluated over a range of chamber pressures and mixture ratios. The specific objectives of the test were: (1) to determine whether stability has been influenced by injection of boundary layer coolant across the cavity entrance, (2) if the injector is stable, to determine the minimum cavity area required to maintain stability, and (3) if the injector is unstable, to determine the effects of entrance geometry and increased area on stability.

Pauckert, R. P.↗

NASA NTRS1980

Cooling of High Pressure Rocket Thrust Chambers with Liquid Oxygen

An experimental program using hydrogen and oxygen as the propellants and supercritical liquid oxygen (LOX) as the coolant was conducted at 4.14 and 8.274 MN/square meters (600 and 1200 psia) chamber pressure. Data on the following are presented: the effect of LOX leaking into the combustion region through small cracks in the chamber wall; and verification of the supercritical oxygen heat transfer correlation developed from heated tube experiments; A total of four thrust chambers with throat diameters of 0.066 m were tested. Of these, three were cyclically tested to 4.14 MN/square meters (600 psia) chamber pressure until a crack developed. One had 23 additional hot cycles accumulated with no apparent metal burning or distress. The fourth chamber was operated at 8.274 MN/square meters (1200 psia) pressure to obtain steady state heat transfer data. Wall temperature measurements confirmed the heat transfer correlation.

Price, H. G.↗

NASA NTRS1990

RS-88 Pad Abort Demonstrator Thrust Chamber Assembly Testing at NASA Marshall Space Flight Center

This paper documents the effort conducted to collect hot-tire dynamic and acoustics environments data during 50,000-lb thrust lox-ethanol hot-fire rocket testing at NASA Marshall Space Flight Center (MSFC) in November-December 2003. This test program was conducted during development testing of the Boeing Rocketdyne RS-88 development engine thrust chamber assembly (TCA) in support of the Orbital Space Plane (OSP) Crew Escape System Propulsion (CESP) Program Pad Abort Demonstrator (PAD). In addition to numerous internal TCA and nozzle measurements, induced acoustics environments data were also collected. Provided here is an overview of test parameters, a discussion of the measurements, test facility systems and test operations, and a quality assessment of the data collected during this test program.

Farr, Rebecca A.↗

NASA NTRS2007

Multiphysics Computational Analysis of a Solid-Core Nuclear Thermal Engine Thrust Chamber

The objective of this effort is to develop an efficient and accurate computational heat transfer methodology to predict thermal, fluid, and hydrogen environments for a hypothetical solid-core, nuclear thermal engine - the Small Engine. In addition, the effects of power profile and hydrogen conversion on heat transfer efficiency and thrust performance were also investigated. The computational methodology is based on an unstructured-grid, pressure-based, all speeds, chemically reacting, computational fluid dynamics platform, while formulations of conjugate heat transfer were implemented to describe the heat transfer from solid to hydrogen inside the solid-core reactor. The computational domain covers the entire thrust chamber so that the afore-mentioned heat transfer effects impact the thrust performance directly. The result shows that the computed core-exit gas temperature, specific impulse, and core pressure drop agree well with those of design data for the Small Engine. Finite-rate chemistry is very important in predicting the proper energy balance as naturally occurring hydrogen decomposition is endothermic. Locally strong hydrogen conversion associated with centralized power profile gives poor heat transfer efficiency and lower thrust performance. On the other hand, uniform hydrogen conversion associated with a more uniform radial power profile achieves higher heat transfer efficiency, and higher thrust performance.

Wang, Ten-See↗

NASA NTRS1993

Design verification test matrix development for the STME thrust chamber assembly

This report presents the results of the test matrix development for design verification at the component level for the National Launch System (NLS) space transportation main engine (STME) thrust chamber assembly (TCA) components including the following: injector, combustion chamber, and nozzle. A systematic approach was used in the development of the minimum recommended TCA matrix resulting in a minimum number of hardware units and a minimum number of hot fire tests.

Dexter, Carol E.↗

NASA NTRS1990

CFD Code Survey for Thrust Chamber Application

In the quest fo find analytical reference codes, responses from a questionnaire are presented which portray the current computational fluid dynamics (CFD) program status and capability at various organizations, characterizing liquid rocket thrust chamber flow fields. Sample cases are identified to examine the ability, operational condition, and accuracy of the codes. To select the best suited programs for accelerated improvements, evaluation criteria are being proposed.

Gross, Klaus W.↗

NASA NTRS1981

Structural analysis of cylindrical thrust chambers, volume 3

A system of three computer programs is described for use in conjunction with the BOPAGE finite element program. The programs are demonstrated by analyzing cumulative plastic deformation in a regeneratively cooled rocket thrust chamber. The codes provide the capability to predict geometric and material nonlinear behavior of cyclically loaded structures without performing a cycle-by-cycle analysis over the life of the structure. The program set consists of a BOPACE restart tape reader routine, and extrapolation program and a plot package.

Pearson, M. L.↗

NASA NTRS2007

Numerical Simulations of Single Flow Element in a Nuclear Thermal Thrust Chamber

The objective of this effort is to develop an efficient and accurate computational methodology to predict both detailed and global thermo-fluid environments of a single now element in a hypothetical solid-core nuclear thermal thrust chamber assembly, Several numerical and multi-physics thermo-fluid models, such as chemical reactions, turbulence, conjugate heat transfer, porosity, and power generation, were incorporated into an unstructured-grid, pressure-based computational fluid dynamics solver. The numerical simulations of a single now element provide a detailed thermo-fluid environment for thermal stress estimation and insight for possible occurrence of mid-section corrosion. In addition, detailed conjugate heat transfer simulations were employed to develop the porosity models for efficient pressure drop and thermal load calculations.

Cheng, Gary↗

NASA NTRS2007

Development of an Efficient CFD Model for Nuclear Thermal Thrust Chamber Assembly Design

The objective of this effort is to develop an efficient and accurate computational methodology to predict both detailed thermo-fluid environments and global characteristics of the internal ballistics for a hypothetical solid-core nuclear thermal thrust chamber assembly (NTTCA). Several numerical and multi-physics thermo-fluid models, such as real fluid, chemically reacting, turbulence, conjugate heat transfer, porosity, and power generation, were incorporated into an unstructured-grid, pressure-based computational fluid dynamics solver as the underlying computational methodology. The numerical simulations of detailed thermo-fluid environment of a single flow element provide a mechanism to estimate the thermal stress and possible occurrence of the mid-section corrosion of the solid core. In addition, the numerical results of the detailed simulation were employed to fine tune the porosity model mimic the pressure drop and thermal load of the coolant flow through a single flow element. The use of the tuned porosity model enables an efficient simulation of the entire NTTCA system, and evaluating its performance during the design cycle.

Cheng, Gary↗

NASA NTRS1977

Hydrogen film cooling of a small hydrogen-oxygen thrust chamber and its effect on erosion rates of various ablative materials

An experimental investigation was conducted to determine what arrangement of film-coolant-injection orifices should be used to decrease the erosion rates of small, high temperature, high pressure ablative thrust chambers without incurring a large penalty in combustion performance. All of the film cooling was supplied through holes in a ring between the outer row of injector elements and the chamber wall. The best arrangement, which had twice the number of holes as there were outer row injection elements, was also the simplest. The performance penalties, presented as a reduction in characteristic exhaust velocity efficiency, were 0.8 and 2.8 percentage points for the 10 and 20 percent cooling flows, respectively, The best film-coolant injector was then used to obtain erosion rates for 19 ablative materials. The throat erosion rate was reduced by a factor of 2.5 with a 10 percent coolant flow. Only the more expensive silica phenolic materials had low enough erosion rates to be considered for use in the nozzle throat. However, some of the cheaper materials might qualify for use in other areas of small nozzles with large throat diameters where the higher erosion rates are more acceptable.

Hannum, N.↗

NASA NTRS1978

Development of sputtering process to deposit stoichiometric zirconia coatings for the inside wall of regeneratively cooled rocket thrust chambers

Thermal barrier coatings of yttria stabilized zirconia and zirconia-ceria mixtures were deposited by RF reactive sputtering. Coatings were 1-2 mils thick, and were deposited on copper cylinders intended to simulate the inner wall of a regeneratively cooled thrust chamber. Coating stoichiometry and adherence were investigated as functions of deposition parameters. Modest deposition rates (approximately 0.15 mil/hr) and subambient sustrate temperatures (-80 C) resulted in nearly stoichiometric coatings which remained adherent through thermal cycles between -196 and 400 C. Coatings deposited at higher rates or substrates temperatures exhibited greater oxygen deficiences, while coatings deposited at lower temperatures were not adherent. Substrate bias resulted in structural changes in the coating and high krypton contents; no clear effect on stoichiometry was observed.

Busch, R.↗

NASA NTRS1972

Space shuttle orbit maneuvering engine reusable thrust chamber

A data dump is presented containing space shuttle orbiter maneuvering engine performance, weight, envelope, and interface pressure requirements for candidate propellant combinations (NTO/MMH, NTO50-50, LOX/MMH, LOX/50-50, LOX/N2H4, LOX/C3H8, and LOX/RP-1) and cooling concepts (regenerative and dump/film). These data are presented parametrically for the thrust, chamber pressure, nozzle expansion ratio, and engine mixture ratio ranges of interest. Also included is information describing sensitivity to system changes; reliability, maintainability and safety; development programs and associated critical technology areas; engine cost comparisons during development and operation; and ecological effects.

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NASA NTRS1986

SSME thrust chamber modeling with Navier Stokes equations

The capability of predicting two-dimensional, compressible and reacting flow in the combustion chamber and nozzle of the Space Shuttle Main Engine (SSME) is demonstrated. A nonorthogonal body fitted coordinate system has been used to represent the combustor and nozzle geometry. The Navier-Stokes equations are solved for the entire thrust chamber with the k-epsilon turbulence model accounting for compressibility and large pressure gradients effects. Results of the computational test cases reveal all expected features of the transonic nozzle flows including location of sonic line, internal shock and boundary layer build-up. Calculated performance parameters such as thrust, flow rate, and specific impulse are also in reasonble agreement with available data. The results show promising potential of solving full Navier-Stokes equations with heat transfer and two-phase combustion in truly comprehensive modeling of rocket engines.

Przekwas, A. J.↗

NASA NTRS1973

Nondestructive evaluation of regeneratively cooled thrust chambers for rocket engines

Review of a development program for nondestructive tests applicable to hardware with electroformed nickel bonds. The four steps of the program include the selection of some most promising nondestructive evaluation methods out of existing methods, the fabrication of test panels for simulation of cooled thrust chamber walls with bonds of various integrities, the evaluation of the test panels by personnel without previous knowledge of the planned bond integrity, and the verification of actual bond strength in the test panels by destructive tests. Ultrasonics, acoustic emission and holography are evaluated in nondestructive testing applications.

Stauffis, C. R.↗

NASA NTRS1975

Space shuttle orbit maneuvering engine reusable thrust chamber program

Analyses and preliminary designs of candidate OME propellant combinations and corresponding engine designs were conducted and evaluated in terms of performance, operating limits, program cost, risk, inherent life and maintainability. For the Rocketdyne recommended and NASA approved propellant combination and cooling concept (NTO/MMH regeneratively cooled engine), a demonstration thrust chamber was designed, fabricated, and experimentally evaluated to define operating characteristics and limits. Alternate fuel (50-50) operating characteristics were also investigated with the demonstration chamber. Adverse operating effects on regenerative cooled operation were evaluated using subscale electrically heated tubes and channels. An investigation of like doublet element characteristics using subscale tests was performed. Full scale 8- and 10-inch diameter like-doublet injectors for the OME were designed, fabricated, and tested. Injector stability was evaluated analytically and experimentally.

Pauckert, R. P.↗

NASA NTRS1990

A life comparison of tube and channel cooling passages for thrust chambers

The life analysis used to compare copper tubes and milled copper channels for rocket engine cooling passages is described. Copper tubes were chosen as a possible replacement for the existing milled copper channel configuration because (1) they offer increased surface area for additional enthalpy extraction; (2) they have ideal pressure vessel characteristics; and (3) the shape of the tube is believed to allow free expansion, thus accommodating the strain resulting from thermal expansion. The analysis was a two-dimensional elastic-plastic comparison, using a finite element method, to illustrate that, under the same thermal and pressure loading, the compliant shape of the tube increases the life of the chamber. The analysis indicates that for a hot-gas-side-wall temperature of 100 F the critical strain decreases from 1.25 percent in the channel to 0.94 percent in the tube. Since the life of rocket thrust chambers is most often limited by cyclic strain or strain range, this decrease corresponds to an expected tube life which is nearly twice the channel life.

Jankovsky, R. S.↗

NASA NTRS1993

STE thrust chamber technology: Main injector technology program and nozzle Advanced Development Program (ADP)

The purpose of the STME Main Injector Program was to enhance the technology base for the large-scale main injector-combustor system of oxygen-hydrogen booster engines in the areas of combustion efficiency, chamber heating rates, and combustion stability. The initial task of the Main Injector Program, focused on analysis and theoretical predictions using existing models, was complemented by the design, fabrication, and test at MSFC of a subscale calorimetric, 40,000-pound thrust class, axisymmetric thrust chamber operating at approximately 2,250 psi and a 7:1 expansion ratio. Test results were used to further define combustion stability bounds, combustion efficiency, and heating rates using a large injector scale similar to the Pratt & Whitney (P&W) STME main injector design configuration including the tangential entry swirl coaxial injection elements. The subscale combustion data was used to verify and refine analytical modeling simulation and extend the database range to guide the design of the large-scale system main injector. The subscale injector design incorporated fuel and oxidizer flow area control features which could be varied; this allowed testing of several design points so that the STME conditions could be bracketed. The subscale injector design also incorporated high-reliability and low-cost fabrication techniques such as a one-piece electrical discharged machined (EDMed) interpropellant plate. Both subscale and large-scale injectors incorporated outer row injector elements with scarfed tip features to allow evaluation of reduced heating rates to the combustion chamber.

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NASA NTRS1979

Development of sputtered techniques for thrust chambers

The 0.152 cm thick sputtered and copper deposits were electron beam welded to wrought copper. Tensile specimens were machined from the weld assemblies and tested at room temperature. Tensile strength approached the strength of wrought material. Elongations up to 25% were measured. Sputtered aluminum was used to fill 0.157 cm wide by 0.127 cm deep grooves in thrust chamber spool piece liners. The liners were closed out by sputtering copper from post and hollow cathodes.

Mullaly, J. R.↗