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An apparatus was designed for use in determining ignition and combustion characteristics of metallic fuel particles and liquid fuel droplets. The design was tailored around the available data acquisition systems and diagnostic instruments. The apparatus consisted of a small windowed combustion bomb, a C02 laser for particle ignition, a thermal imaging camera for recording the ignition process, a helium-neon laser and a 1024 element linear diode array sampled at 1Mhz for measuring forward scattered light and determining particle size (Fraunhofer diffraction). All major components were individually utilized. However, equipment limitations prevented actual use of the system for measurement of the burning rates of particles. Recommendations are made which would permit the original objectives to be met, http://archive.org/details/anpparatusformea1094526882, Lieutenant, United States Navy, Approved for public release; distribution is unlimited.

The shock structure at the tip of a transonic compressor rotor was investigated on a water table. A four bladed cascade model was used and the wave pattern was examined at variable incidence, flow turning angles and back-pressures. Froude numbers, (equivalent to Mach numbers in the analogous two dimensional gas flow), in the range 1.6 to 1.74 resulted in an oblique shock between the blade passages starting from the leading edge pressure side of the blading. Qualitative agreement of the shock structure with earlier tests using the same blading in a transonic blow-down tunnel was observed, leading to the conclusion that the shock present in the compressor would be oblique and not normal as was previously assumed., http://archive.org/details/transoniccompres1094522158, Approved for public release; distribution is unlimited.

Large cyclical pressure excursions from stator-averaged values are observed on the suction side of a rotor blade at the case wall in a multistage compressor as the rotor moves relative to the stator. The pressure changes correlate well with stator relative position, occur in the passage away from the blade, and are only slightly modified by tip clearance changes. Close to the entry and exit of the tip gap, however, the pressures remain stable. The pressures near the gap are only slightly affected by stator proximity and primarily vary with tip gap height.

An apparatus was designed for use in determining ignition and combustion characteristics of metallic fuel particles and liquid fuel droplets. The design was tailored around the available data acquisition systems and diagnostic instruments. The apparatus consisted of a small windowed combustion bomb, a C02 laser for particle ignition, a thermal imaging camera for recording the ignition process, a helium-neon laser and a 1024 element linear diode array sampled at 1Mhz for measuring forward scattered light and determining particle size (Fraunhofer diffraction). All major components were individually utilized. However, equipment limitations prevented actual use of the system for measurement of the burning rates of particles. Recommendations are made which would permit the original objectives to be met

Detailed two-component laser Doppler velocimeter (LDV) measurements of the flow through a controlled-diffusion (CD) compressor cascade at a Reynolds number of about 700,000 and at a low Mach number are reported in this article. A very high-incidence angle (8 deg above design) was considered throughout this investigation, which included the full experimental characterization of the turbulence Held. The LDV measurements were fully automated and were all taken in coincidence mode, thus turbulent flow correlations could be determined. Most significant was the measurement of the distortion of the inlet freestream turbulence upstream of the blade leading edges. Such information is important in assessing viscous codes which incorporate transport equations to describe the turbulence within the flowfield. The laminar leading-edge separation bubble, which reattached turbulent, was enlarged on the suction surface of the blade. Consistent with measurements at lower incidence angles, the suction surface boundary layer remained attached over the rear part of the blade. The pressure side boundary layer initially showed little or no growth, however, it finally developed into a profile similar to a wall jet. The wake profiles showed significant asymmetry due to the high loading on the blades at the increased incidence angle.

This paper describes the effect of tip clearance changes on the pressure at the case wall of a second stage rotor. Wall shear distributions under the rotor tip are also presented. The results show low pressure areas extending along the rotor suction side but lying away from the blade. Pressure contours indicate the tangential loading at the tip is lower than predicted by two dimensional calculations, however, the predicted loading is observed between the lowest pressure’s path in the passage and the blade pressure side. The results suggest a viscous or shearing layer, due to blade-to-wall relative motion, is generated on the blade side of the tip gap which modifies the inviscid relative flow field and produces an unloading on the blade tip.

An initial experimental investigation was conducted to examine the feasibility of NO x emission control using catalytic reduction techniques in the jet engine test cell environment. A modified T-63 gas turbine combustor and an augmentor tube, 21 feet in length and containing a perlite catalyst, were used as a gas generator and catalytic reduction system. Four data runs were made. Three runs were completed without the catalyst installed. Temperature and velocity profile measurements were obtained in order to calculate augmentation ratios for different engine fuel to air ratios. NOx, CO, and unburned hydrocarbon concentrations in the exhaust were measured to provide a baseline for further tests. A fourth data run was made with the perlite catalyst installed in the augmentor tube. A 64 percent NOx reduction was observed, however, the large pressure drop across the catalytic bed deemed the current configuration impractical Recommendations for alternative configurations are presented. The results of the investigation have proven that further study is warranted.

http://archive.org/details/aninvestigationi1094537988, Lieutenant Commander, Unit ed States Navy, Approved for public release; distribution is unlimited.

Development of the subsonic cascade wind tunnel facility required determination of the two-dimensionality and periodicity of the airflow in the test section with test cascade installed. Data acquisition procedures were developed, and data were recorded for two facility configurations. The flow was shown to be unsatisfactory at a diffusion factor of approximately 0.5S and aspect ratio 1.25, and to be acceptably two-dimensional and periodic at a diffusion factor of approximately 0.39 and aspect ratio 1.95.

An existing code for calculating axial turbine performance using multiple stream surfaces was modified and made to run on the equivalent of an HP- 1000 computer system. Calculations were made for the geometry of a 485 horsepower dual-discharge air-drive turbine for both on and off -design conditions. The results were compared with available data obtained at off -design speeds. Agreement of the flow rate and horsepower to within 51 was obtained.

The shock structure at the tip of a transonic compressor rotor was investigated on a water table. A four bladed cascade model was used and the wave pattern was examined at variable incidence, flow turning angles and back-pressures. Froude numbers, (equivalent to Mach numbers in the analogous two dimensional gas flow), in the range 1.6 to 1.74 resulted in an oblique shock between the blade passages starting from the leading edge pressure side of the blading. Qualitative agreement of the shock structure with earlier tests using the same blading in a transonic blow-down tunnel was observed, leading to the conclusion that the shock present in the compressor would be oblique and not normal as was previously assumed.

Tests are reported in which air incidence angle was varied to a cascade of 20 blades, 5 inches in chord with aspect ratio of 2.0 and solidity of 1.67. Preliminary blade element performance data were obtained using pneumatic probe surveys and surface pressures were also measureed Results of preparatory tests of a similar cascade of 15 C-series blades at a solidity of 1.28 are also included. Whereas the flow in the C-series blading cascade was always acceptable, the flow in the DCA blading cascade was not acceptable at negative incidence angles. A modification of the cascade inlet guide vanes is recommended which will guarantee periodic conditions and serve to generate more nearly uniform inlet flow at all test conditions.

The design and construction of a small, transonic, cascade wind tunnel are reported. The tunnel is of the intermittent blow-down type, and the initial test cascade models the flow at the rotor blade tips of a single stage transonic compressor at a relative Mach number of 1.U-. General programs were developed for the design of round-torectangular transition sections and for the computation of supersonic nozzle contours. Complete machine drawings for the important components of the facility are included.

A detailed two-component LDV mapping of the flow through a controlled diffusion compressor cascade at low Mach number (∼0.25) and Reynolds number of about 7×105, at three inlet air angles from design to near stall, is reported. It was found that the suction side boundary layer reattached turbulent after a laminar separation bubble, and remained attached to the trailing edge even at the highest incidence, at which losses were 3 to 4 times the minimum value for the geometry. Boundary layer thickness increased to fill 20% of the blade passage at the highest incidence. Results for pressure-side boundary layer and near wake are also summarized. Information sufficient to allow preliminary assessment of viscous codes is tabulated.

The mid-span section of a previously reported controlled-diffusion compressor stator has been experimentally evaluated in cascade. Measurements are taken over a range of incidence angles for blade chord Reynolds numbers from 470,000 to 690,000. Blade chord length is 12.7 cm, aspect ratio is 2.0, and solidity is 1.67. Measurements include conventional cascade performance parameters as well as blade surface pressures. Computations are made for the inviscid flow field, surface boundary layers, and loss for several of the blade inlet angle conditions, are compared against corresponding data.

A large scale (2 in.-wide test vanes at 50-in. inlet diameter), low speed (100-200 ft/sec) steady flow radial cascade wind tunnel for diffuser studies was designed, built and tested. The apparatus was shown to provide flow angles (from radial) in the range of 58-72 deg with suitable spanwise profiles. The novel flow angle control mechanism was shown to work, but measured flow angles were somewhat smaller than expected. The mechanism for controlling case-wall boundary layer profile did not behave as predicted. Attempts to predict the generated flow both analytically and with computational codes are compared with initial measurements. The low speed apparatus will be used to obtain detailed data for diffuser design and analysis code verification, and to provide experience toward the design of a high speed device.

Exploratory experiments with an air-breathing propulsion device which uses detonative combustion to achieve high cycle pressures, are reported. In the first phase, detonation in a chamber containing ethylene and air was initiated by a detonation produced in a small tube containing ethylene and oxygen. An automobile spark system was used for ignition. In the second phase, a repetitive detonation process was sustained in a naturally-aspirated chamber at a frequency of 25Hz. Projections based on the present test results suggest that frequencies of 150Hz and specific impulse in the range 1000-1400 sec might be obtained in a practical engine.

Exploratory experiments with an air-breathing propulsion device which uses detonative combustion to achieve high cycle pressures, are reported. In the first phase, detonation in a chamber containing ethylene and air was initiated by a detonation produced in a small tube containing ethylene and oxygen. An automobile spark system was used for ignition. In the second phase, a repetitive detonation process was sustained in a naturally-aspirated chamber at a frequency of 25Hz. Projections based on the present test results suggest that frequencies of 150Hz and specific impulse in the range 1000-1400 sec might be obtained in a practical engine.

The purpose of this brief presentation is to report work in progress at the Turbopropulsion Laboratory at the Naval Postgraduate School to determine experimentally the behavior of the flow in a transonic compressor stage. The intent is to develop an input into the design method which accounts , in some way, for the three dimensional and periodic nature of the flow. While this has been the goal of continuing research efforts, no significant change in the basic design model has been reported. The essential three dimensional nature of transonic flow is acknowledged, but no design method is three dimensional. Within the pseudo two dimensional analytical framework, secondary flows are known to become more significant as the blade loading is increased, and wall effects become important as the passage depth is decreased. While the effects are qualitatively known to be significant, the understanding that comes from favorable comparisons of observation. with analyses is lacking . First, definitive measurements from different stage designs are needed to define the relative importance of different effects and thus provide the basis for a representative analytical design model. In order to achieve higher performance from a new design without a costly empirical development process, a design model which more nearly represents the actual flow In the blade rows is essential.

An existing code for calculating axial turbine performance using multiple stream surfaces was modified and made to run on the equivalent of an HP- 1000 computer system. Calculations were made for the geometry of a 485 horsepower dual-discharge air-drive turbine for both on and off -design conditions. The results were compared with available data obtained at off -design speeds. Agreement of the flow rate and horsepower to within 51 was obtained., http://archive.org/details/computerevaluati00ciro, Lieutenant, United States Navy, Approved for public release; distribution is unlimited.

Development of the subsonic cascade wind tunnel facility required determination of the two-dimensionality and periodicity of the airflow in the test section with test cascade installed. Data acquisition procedures were developed, and data were recorded for two facility configurations. The flow was shown to be unsatisfactory at a diffusion factor of approximately 0.5S and aspect ratio 1.25, and to be acceptably two-dimensional and periodic at a diffusion factor of approximately 0.39 and aspect ratio 1.95., http://archive.org/details/evaluationofsubs1094517580, Lieutenant Commander, United States Navy, Approved for public release; distribution is unlimited.

The design and construction of a small, transonic, cascade wind tunnel are reported. The tunnel is of the intermittent blow-down type, and the initial test cascade models the flow at the rotor blade tips of a single stage transonic compressor at a relative Mach number of 1.U-. General programs were developed for the design of round-torectangular transition sections and for the computation of supersonic nozzle contours. Complete machine drawings for the important components of the facility are included., http://archive.org/details/cascadewindtunne1094518365, Lieutenant, United States Navy, Approved for public release; distribution is unlimited.

An overview, presentation and discussion in a 26/27th 6/27th March 1985 workshop held at the Naval Postgraduate School, "to provide the input necessary for an assessment of the prospects for the useful exploitation of wave rotors in gas turbine engines", are documented., Naval Research (Mechanics Division) 800 N. Quincy Street, Arlington VA 2217, N0001485WR24208, N0001985WR51147, Approved for public release; distribution unlimited.

Tests are reported in which air incidence angle was varied to a cascade of 20 blades, 5 inches in chord with aspect ratio of 2.0 and solidity of 1.67. Preliminary blade element performance data were obtained using pneumatic probe surveys and surface pressures were also measureed Results of preparatory tests of a similar cascade of 15 C-series blades at a solidity of 1.28 are also included. Whereas the flow in the C-series blading cascade was always acceptable, the flow in the DCA blading cascade was not acceptable at negative incidence angles. A modification of the cascade inlet guide vanes is recommended which will guarantee periodic conditions and serve to generate more nearly uniform inlet flow at all test conditions., http://archive.org/details/subsoniccascadew1094520690, Lieutenant, United States Navy, Approved for public release; distribution is unlimited.

An initial experimental investigation was conducted to examine the feasibility of NO x emission control using catalytic reduction techniques in the jet engine test cell environment. A modified T-63 gas turbine combustor and an augmentor tube, 21 feet in length and containing a perlite catalyst, were used as a gas generator and catalytic reduction system. Four data runs were made. Three runs were completed without the catalyst installed. Temperature and velocity profile measurements were obtained in order to calculate augmentation ratios for different engine fuel to air ratios. NOx, CO, and unburned hydrocarbon concentrations in the exhaust were measured to provide a baseline for further tests. A fourth data run was made with the perlite catalyst installed in the augmentor tube. A 64 percent NOx reduction was observed, however, the large pressure drop across the catalytic bed deemed the current configuration impractical Recommendations for alternative configurations are presented. The results of the investigation have proven that further study is warranted., http://archive.org/details/anexperimentalin1094543846, Lieutenant Commander, United States Navy, Approved for public release; distribution is unlimited.