ALBERT

Description: Two nonintrusive flowfield diagnostics based on spectrally-resolved elastic (Rayleigh) and inelastic (Raman) laser light scattering were developed for obtaining local flowfield measurements in low-thrust gaseous H2/O2 rocket engines. The objective is to provide an improved understanding of phenomena occurring in small chemical rockets in order to facilitate the development and validation of advanced computational fluid dynamics (CFD) models for analyzing engine performance. The laser Raman scattering diagnostic was developed to measure major polyatomic species number densities and rotational temperatures in the high-density flowfield region extending from the injector through the chamber throat. Initial application of the Raman scattering diagnostic provided O2 number density and rotational temperature measurements in the exit plane of a low area-ratio nozzle and in the combustion chamber of a two-dimensional, optically-accessible rocket engine. In the low-density nozzle exit plane region where the Raman signal is too weak, a Doppler-resolved laser Rayleigh scattering diagnostic was developed to obtain axial and radial mean gas velocities, and in certain cases, H2O translational temperature and number density. The results from these measurements were compared with theoretical predictions from the RPLUS CFD code for analyzing rocket engine performance. Initial conclusions indicate that a detailed and rigorous modeling of the injector is required in order to make direct comparisons between laser diagnostic measurements and CFD predictions at the local level.

Description: The NASA Glenn Research Center is working with the aeronautics industry to develop highly fuel-efficient and environmentally friendly gas turbine combustor technology. This effort includes testing new hardware designs at conditions that simulate the high-temperature, high-pressure environment expected in the next-generation of high-performance engines. Glenn has the only facilities in which such tests can be performed. One aspect of these tests is the use of nonintrusive optical and laser diagnostics to measure combustion species concentration, fuel/air ratio, fuel drop size, and velocity, and to visualize the fuel injector spray pattern and some combustion species distributions. These data not only help designers to determine the efficacy of specific designs, but provide a database for computer modelers and enhance our understanding of the many processes that take place within a combustor. Until recently, we lacked one critical capability, the ability to measure temperature. This article summarizes our latest developments in that area. Recently, we demonstrated the first-ever use of spontaneous Raman scattering to measure combustion temperatures within the Advanced Subsonics Combustion Rig (ASCR) sector rig. We also established the highest rig pressure ever achieved for a continuous-flow combustor facility, 54.4 bar. The ASCR facility can provide operating pressures from 1 to 60 bar (60 atm). This photograph shows the Raman system setup next to the ASCR rig. The test was performed using a NASA-concept fuel injector and Jet-A fuel over a range of air inlet temperatures, pressures, and fuel/air ratios.

Description: The NASA Glenn Research Center and the aerospace industry are designing and testing low-emission combustor concepts to build the next generation of cleaner, more fuel efficient aircraft powerplants. These combustors will operate at much higher inlet temperatures and at pressures that are up to 3 to 5 times greater than combustors in the current fleet. From a test and analysis viewpoint, there is an increasing need for measurements from these combustors that are nonintrusive, simultaneous, multipoint, and more quantitative. Glenn researchers have developed several unique test facilities (refs. 1 and 2) that allow, for the first time, optical interrogation of combustor flow fields, including subcomponent performance, at pressures ranging from 1 to 60 bar (1 to 60 atm). Experiments conducted at Glenn are the first application of a visible laser-pumped, one-dimensional, spontaneous Raman-scattering technique to analyze the flow in a high-pressure, advanced-concept fuel injector at pressures thus far reaching 12 bar (12 atm). This technique offers a complementary method to the existing two- and three-dimensional imaging methods used, such as planar laser-induced fluorescence. Raman measurements benefit from the fact that the signal from each species is a linear function of its density, and the relative densities of all major species can be acquired simultaneously with good precision. The Raman method has the added potential to calibrate multidimensional measurements by providing an independent measurement of species number-densities at known points within the planar laser-induced fluorescence images. The visible Raman method is similar to an ultraviolet-Raman technique first tried in the same test facility (ref. 3). However, the visible method did not suffer from the ultraviolet technique's fuel-born polycyclic aromatic hydrocarbon fluorescence interferences.

Description: Researchers at the NASA Glenn Research Center, in collaboration with industry, are reducing gas turbine engine emissions by studying visually the air-fuel interactions and combustion processes in combustors. This is especially critical for next generation engines that, in order to be more fuel-efficient, operate at higher temperatures and pressures than the current fleet engines. Optically based experiments were conducted in support of the Ultra-Efficient Engine Technology program in Glenn's unique, world-class, advanced subsonic combustion rig (ASCR) facility. The ASCR can supply air and jet fuel at the flow rates, temperatures, and pressures that simulate the conditions expected in the combustors of high-performance, civilian aircraft engines. In addition, this facility is large enough to support true sectors ("pie" slices of a full annular combustor). Sectors enable one to test true shapes rather than rectangular approximations of the actual hardware. Therefore, there is no compromise to actual engine geometry. A schematic drawing of the sector test stand is shown. The test hardware is mounted just upstream of the instrumentation section. The test stand can accommodate hardware up to 0.76-m diameter by 1.2-m long; thus sectors or small full annular combustors can be examined in this facility. Planar (two-dimensional) imaging using laser-induced fluorescence and Mie scattering, chemiluminescence, and video imagery were obtained for a variety of engine cycle conditions. The hardware tested was a double annular sector (two adjacent fuel injectors aligned radially) representing approximately 15 of a full annular combustor. An example of the two-dimensional data obtained for this configuration is also shown. The fluorescence data show the location of fuel and hydroxyl radical (OH) along the centerline of the fuel injectors. The chemiluminescence data show C2 within the total observable volume. The top row of this figure shows images obtained at an engine low-power condition, and the bottom row shows data from a higher power operating point. The data show distinctly the differences in flame structure between low-power and high-power engine conditions, in both location and amount of species produced (OH, C2) or consumed (fuel). The unique capability of the facility coupled with its optical accessibility helps to eliminate the need for high-pressure performance extrapolations. Tests such as described here have been used successfully to assess the performance of fuel-injection concepts and to modify those designs, if needed.

Description: The emphasis of combustion research efforts at NASA Glenn Research Center (GRC) is on collaborating with industry to design and test gas-turbine combustors and subcomponents for both sub- and supersonic applications. These next-generation aircraft combustors are required to meet strict international environmental restrictions limiting emissions. To meet these goals, innovative combustor concepts require operation at temperatures and pressures far exceeding those of cur-rent designs. New and innovative diagnostic tools are necessary to characterize these flow streams since existing methods are inadequate. The combustion diagnostics team at GRC has implemented a suite of highly sensitive, nonintrusive optical imaging methods to diagnose the flowfields of these new engine concepts. By using optically accessible combustors and flametubes, imaging of fuel and intermediate combustion species via planar laser-induced fluorescence (PLIF) at realistic pressures are now possible. Direct imaging of the fuel injection process through both planar Mie scattering and PLIF methods is also performed. Additionally, a novel combination of planar fuel fluorescence imaging and computational analysis allows a 3-D examination of the flowfield, resulting in spatially and temporally resolved fuel/air volume distribution maps. These maps provide detailed insight into the fuel injection process at actual conditions, thereby greatly enhancing the evaluation of fuel injector performance and other combustion phenomena. Stable species such as CO2, O2, N2O. and hydrocarbons are also investigated by a newly demonstrated 1-D, spontaneous Raman spectroscopic method. This visible wavelength Raman technique allows the acquisition of quantitative. stable species concentration measurements from the flow.

Description: A decomposing solid thruster concept, which creates a more benign thermal and chemical environment than solid propellant combustion, while maintaining, performance similar to solid combustion, is described. A Micro-Electro-Mechanical (MEMS) thruster concept with diode laser and fiber-optic initiation is proposed, and thruster components fabricated with MEMS technology are presented. A high nitrogen content solid gas generator compound is evaluated and tested in a conventional axisymmetric thrust chamber with nozzle throat area ratio of 100. Results show incomplete decomposition of this compound in both low pressure (1 kPa) and high pressure (1 MPa) environments, with decomposition of up to 80% of the original mass. Chamber pressures of 1.1 MPa were obtained, with maximum calculated thrust of approximately 2.7 N. Resistively heated wires and resistively heated walls were used to initiate decomposition. Initiation tests using available lasers were unsuccessful, but infrared spectra of the compound show that the laser initiation tests used inappropriate wavelengths for optimal propellant absorption. Optimal wavelengths for laser ignition were identified. Data presented are from tests currently in progress. Alternative solid gas generator compounds are being evaluated for future tests.

Description: Localized velocity, temperature, and species concentration measurements in rocket flow fields are needed to evaluate predictive computational fluid dynamics (CFD) codes and identify causes of poor rocket performance. Velocity, temperature, and total number density information have been successfully extracted from spectrally resolved Rayleigh scattering in the plume of small hydrogen/oxygen rockets. Light from a narrow band laser is scattered from the moving molecules with a Doppler shifted frequency. Two components of the velocity can be extracted by observing the scattered light from two directions. Thermal broadening of the scattered light provides a measure of the temperature, while the integrated scattering intensity is proportional to the number density. Spontaneous Raman scattering has been used to measure temperature and species concentration in similar plumes. Light from a dye laser is scattered by molecules in the rocket plume. Raman spectra scattered from major species are resolved by observing the inelastically scattered light with linear array mounted to a spectrometer. Temperature and oxygen concentrations have been extracted by fitting a model function to the measured Raman spectrum. Results of measurements on small rockets mounted inside a high altitude chamber using both diagnostic techniques are reported.

Description: Spontaneous vibrational Raman scattering was used to measure temperature in an aviation combustor sector burning jet fuel. The inlet temperature ranged from 670 K (750 F) to 756 K (900 F) and pressures from 13 to 55 bar. With the exception of a discrepancy that we attribute to soot, good agreement was seen between the Raman-derived temperatures and the theoretical temperatures calculated from the inlet conditions. The technique used to obtain the temperature uses the relationship between the N2 anti-Stokes and Stokes signals, within a given Raman spectrum. The test was performed using a NASA-concept fuel injector and Jet-A fuel over a range of fuel/air ratios. This work represents the first such measurements in a high-pressure, research aero-combustor facility.