Engineering topics

Anderson, W. E.

Publications and source records attributed to Anderson, W. E..

17 records

NASA NTRS2005

Subscale Test Methods for Combustion Devices

Stated goals for long-life LRE s have been between 100 and 500 cycles: 1) Inherent technical difficulty of accurately defining the transient and steady state thermochemical environments and structural response (strain); 2) Limited statistical basis on failure mechanisms and effects of design and operational variability; and 3) Very high test costs and budget-driven need to protect test hardware (aversion to test-to-failure). Ambitious goals will require development of new databases: a) Advanced materials, e.g., tailored composites with virtually unlimited property variations; b) Innovative functional designs to exploit full capabilities of advanced materials; and c) Different cycles/operations. Subscale testing is one way to address technical and budget challenges: 1) Prototype subscale combustors exposed to controlled simulated conditions; 2) Complementary to conventional laboratory specimen database development; 3) Instrumented with sensors to measure thermostructural response; and 4) Coupled with analysis

Anderson, W. E.↗

NASA NTRS1999

Development of a Pressure-Fed Rocket Engine Using Hydrogen Peroxide and JP-8

A pressure-fed rocket engine using hydrogen peroxide and JP-8 was designed and demonstrated for upper stage space transportation applications. The engine utilizes silver plated screen catalyst to decompose 85% hydrogen peroxide (by weight). The decomposed high-temperature gas causes the JP-8 to auto-ignite and undergo chemical reaction. An ablative chamber using silica phenolic was used as the combustion chamber. The present effort includes catalyst development, injector evaluation, and chamber char and erosion characterization. Wagonwheel-type distribution and support plates were used in the catalyst bed design to reduce pressure drop, to increase active area, and to avoid flow channeling. Various fuel injection and mixing approaches were investigated and the combustion efficiency was in the range from 89 to 98%, depending on injector design and operating conditions. Fuel film cooling, up to 40% of the total fuel flow, was found to be ineffective with the current engine configuration. A ring injector was selected because of the satisfactory performance and its low cost. The char and erosion rates were determined through hot-fire tests. The char depth was found to be a function of axial distance, which was related to the progress of chemical reaction. Correlation of the char and erosion rates was developed as a design database for future applications,

Wu, P.-K.↗

NASA NTRS1999

Catalyst Development for Hydrogen Peroxide Rocket Engines

The development of various catalysts of hydrogen peroxide was conducted for the applications of liquid rocket engines. The catalyst development includes silver screen technology, solid catalyst technology, and homogeneous catalyst technology. The silver screen technology development was performed with 85% (by weight) hydrogen peroxide. The results of this investigation were used as the basis for the catalyst design of a pressure-fed liquid-fueled upper stage engine. Both silver-plated nickel 200 screens and pure silver screens were used as the active metal catalyst during the investigation, The data indicate that a high decomposition efficiency (greater than 90%) of 85% hydrogen peroxide can be achieved at a bed loading of 0.5 lbm/sq in/sec with both pure silver and silver plated screens. Samarium oxide coating, however, was found to retard the decomposition process and the catalyst bed was flooded at lower bed loading. A throughput of 200 lbm of hydrogen peroxide (1000 second run time) was tested to evaluate the catalyst aging issue and performance degradation was observed starting at approximately 400 seconds. Catalyst beds of 3.5 inch in diameter was fabricated using the same configuration for a 1,000-lbf rocket engine. High decomposition efficiency was obtained with a low pressure drop across the bed. Solid catalyst using precious metal was also developed for the decomposition of hydrogen peroxide from 85% to 98% by weight. Preliminary results show that the catalyst has a strong reactivity even after 15 minutes of peroxide decomposition. The development effort also includes the homogeneous catalyst technology. Various non-toxic catalysts were evaluated with 98% peroxide and hydrocarbon fuels. The results of open cup drop tests indicate an ignition delay around 11 ms.

Morlan, P. W.↗

NASA NTRS1999

Upper-Stage Flight Experiment

For propulsion applications that require that the propellants are storable for long periods, have a high density impulse, and are environmentally clean and non-toxic, the best choice is a combination of high-concentration hydrogen peroxide (High Test Peroxide, or HTP) and a liquid hydrocarbon (LHC) fuel. The HTP/LHC combination is suitable for low-cost launch vehicles, space taxi and space maneuvering vehicles, and kick stages. Orbital Sciences Corporation is under contract with the NASA Marshall Space Flight Center in cooperation with the Air Force Research Lab to design, develop and demonstrate a new low-cost liquid upper stage based on HTP and JP-8. The Upper Stage Flight Experiment (USFE) focuses on key technologies necessary to demonstrate the operation of an inherently simple propulsion system with an innovative, state-of-the-art structure. Two key low-cost vehicle elements will be demonstrated - a 10,000 lbf thrust engine and an integrated composite tank structure. The suborbital flight test of the USFE is scheduled for 2001. Preceding the flight tests are two major series of ground tests at NASA Stennis Space Center and a subscale tank development program to identify compatible composite materials and to verify their compatibility over long periods of time. The ground tests include a thrust chamber development test series and an integrated stage test. This paper summarizes the results from the first phase of the thrust chamber development tests and the results to date from the tank material compatibility tests. Engine and tank configurations that meet the goals of the program are described.

Anderson, W. E.↗

NASA NTRS1997

Rocket Ejector Studies for Application to RBCC Engines: An Integrated Experimental/CFD Approach

Recent interest in low cost, reliable access to space has generated increased interest in advanced technology approaches to space transportation systems. A key to the success of such programs lies in the development of advanced propulsion systems capable of achieving the performance and operations goals required for the next generation of space vehicles. One extremely promising approach involves the combination of rocket and air- breathing engines into a rocket-based combined-cycle engine (RBCC). A key element of that engine is the rocket ejector which is utilized in the zero to Mach two operating regime. Studies of RBCC engine concepts are not new and studies dating back thirty years are well documented in the literature. However, studies focused on the rocket ejector mode of the RBCC cycle are lacking. The present investigation utilizes an integrated experimental and computation fluid dynamics (CFD) approach to examine critical rocket ejector performance issues. In particular, the development of a predictive methodology capable of performance prediction is a key objective in order to analyze thermal choking and its control, primary/secondary pressure matching considerations, and effects of nozzle expansion ratio. To achieve this objective, the present study emphasizes obtaining new data using advanced optical diagnostics such as Raman spectroscopy and CFD techniques to investigate mixing in the rocket ejector mode. A new research facility for the study of the rocket ejector mode is described along with the diagnostic approaches to be used. The CFD modeling approach is also described along with preliminary CFD predictions obtained to date.

Pal, S.↗

NASA NTRS1994

Analytical and experimental studies of impinging liquid jets

Impinging injectors are a common type of injector used in liquid propellant rocket engines and are typically used in engines where both propellants are injected as a liquid, e.g., engines using LOX/hydrocarbon and storable propellant combinations. The present research program is focused on providing the requisite fundamental understanding associated with impinging jet injectors for the development of an advanced a priori combustion stability design analysis capability. To date, a systematic study of the atomization characteristics of impinging liquid jets under cold-flow conditions have been completed. Effects of orifice diameter, impingement angle, pre-impingement length, orifice length-to-diameter ratio, fabrication procedure, jet flow condition and jet velocity under steady and oscillating, and atmospheric- and high-pressure environments have been investigated. Results of these experimental studies have been compared to current models of sheet breakup and drop formation. In addition, the research findings have been scrutinized to provide a fundamental explanation for a proven empirical correlation used in the design of stable impinging injector-based rocket engines.

Ryan, H. M.↗

NASA NTRS1993

Spray formation processes of impinging jet injectors

A study examining impinging liquid jets has been underway to determine physical mechanisms responsible for combustion instabilities in liquid bi-propellant rocket engines. Primary atomization has been identified as an important process. Measurements of atomization length, wave structure, and drop size and velocity distribution were made under various ambient conditions. Test parameters included geometric effects and flow effects. It was observed that pre-impingement jet conditions, specifically whether they were laminar or turbulent, had the major effect on primary atomization. Comparison of the measurements with results from a two dimensional linear aerodynamic stability model of a thinning, viscous sheet were made. Measured turbulent impinging jet characteristics were contrary to model predictions; the structure of waves generated near the point of jet impingement were dependent primarily on jet diameter and independent of jet velocity. It has been postulated that these impact waves are related to pressure and momentum fluctuations near the impingement region and control the eventual disintegration of the liquid sheet into ligaments. Examination of the temporal characteristics of primary atomization (ligament shedding frequency) strongly suggests that the periodic nature of primary atomization is a key process in combustion instability.

Anderson, W. E.↗

NASA NTRS1993

Atomization characteristics of impinging liquid jets

A systematic study is presented of the atomization of impinging liquid jets investigating the effects of jet conditions (laminar versus turbulent), orifice diameter, impingement angle and jet velocity. Results are compared to current theories in terms of sheet breakup length, sheet shape and drop size. Experiments contrasting laminar and turbulent jet conditions clearly show that the jet conditions have a dramatic effect on the atomization process.

Ryan, H. M.↗

NASA NTRS1992

Fundamental studies of impinging liquid jets

Atomization caused by impinging two liquid jets at an angle was studied experimentally and analytically. Measurements of drop size, sheet length and width at breakup, and apparent wave structures on the liquid sheet surface were made for different flow conditions and different geometries. The breakup of the attenuating liquid sheet into ligaments was modeled by linear stability theory. Breakup lengths were predicted from the fastest growing wave on the sheet surface, and drop diameters were calculated based on surface tension-controlled breakup of the ligaments.

Anderson, W. E.↗

NASA NTRS1973

Experimental investigation of a large-scale, two-dimensional, mixed-compression inlet system: Internal performance and drag at transonic conditions, free stream Mach equals 0.6 to 1.28

A large scale, variable-geometry inlet system with a design Mach number of 3.0 was tested at Mach numbers from 0.6 to 1.28. Variable features for off-design operation are an adjustable-height ramp system and a translating cowl. Experimental results are presented for transonic ramp and cowl positions showing the effect of throat boundary layer bleed and vortex generators on engine-face performance. Detailed pressure and force-balance data are used to evaluate transonic drag characteristics.

Wong, N. D.↗

NASA NTRS1971