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RESEARCH PAPERS

J. Eng. Gas Turbines Power. 1985;107(4):808-814. doi:10.1115/1.3239815.

A choice between a derivative engine or a new engine for either an existing aircraft or a new aircraft is becoming an increasingly frequent one. Although some factors which govern such a choice are necessarily subjective, there are several objective factors which should be considered. Using summaries of recent or current cases as examples, an examination of the many dimensions of capability, cost, and risk identifies twelve of these objective factors, many of which are often not subjected to rational analysis.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):815-820. doi:10.1115/1.3239816.

Evolving aircraft and engine technologies, as well as advancements in avionics, have resulted in a new generation of fighter aircraft. In order for an air force to remain survivable in such an environment it must face the necessity of either acquiring new aircraft or upgrading its fleet. Today, acquisition of new aircraft may not be an economically or politically viable solution. The newest aircraft, those with a thrust-to-weight ratio exceeding 1 to 1, generally exceed $ 20 million, exclusive of support costs. Politically, sale of such state-of-the-art aircraft to friendly countries may not be possible given the current FX aircraft export policy, and the fiscal posture of the country. However, there is an alternative to new aircraft acquisition, and that is functional modernization: the modernization of potentially viable systems through the implementation of new and available technologies. The Boeing Military Airplane Company and Pratt & Whitney have teamed in the newest aircraft functional modernization proposal, the F-4 Super Phantom. The full modernization package includes a new centerline conformal fuel tank, new digital avionics, and re-engining with new PW1120 engines. This paper examines the initial feasibility study conducted by Pratt & Whitney and Boeing to determine the feasibility of a functionally modernized F-4.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):821-827. doi:10.1115/1.3239817.

The technology of high-pressure air or hot-gas impingement from stationary shroud supplementary nozzles onto radial outflow compressors and radial inflow turbines to permit rapid gas turbine starting or power boosting is discussed. Data are presented on the equivalent turbine component performance for convergent/divergent shroud impingement nozzles, which reveal the sensitivity of nozzle velocity coefficient with Mach number and turbine efficiency with impingement nozzle admission arc. Compressor and turbine matching is addressed in the transient turbine start mode with the possibility of operating these components in braking or reverse flow regimes when impingement flow rates exceed design.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):828-832. doi:10.1115/1.3239818.

Severe aerodynamic interaction between the fan core stream section and the high-pressure compressor of a three-shaft low-bypass-ratio engine is described. At high fan running lines a heavy single-cell rotating stall was found in the fan core stream even at high aerodynamic speeds between 90–98% . The rotating circumferential distortion with 180–200 deg sector angle is swallowed by the intermediate pressure compressor but erodes the high-pressure compressor surge margin by about 22% , leading to steady-state surges. A remotely mounted transducer in a specific arrangement was used successfully for measurements in the hot environment behind intermediate and high-pressure compressor using a so-called “long-line” system with a closed end at the downstream pipe.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):833-837. doi:10.1115/1.3239819.

Flight and propulsion controls can effectively be integrated to provide operational benefits to the weapon system which cannot be achieved with either system acting independently. The key factor is the synergistic effect of digital computers and associated high-speed data links, engine condition monitoring, diagnostics, and shared power systems (hydraulic and electrical), thus providing potential improvements in reliability, life cycle cost, weight, and maintenance actions. This technical paper will address the integration concepts required for the airframe and the propulsion control configurations (for consideration in the design stages) to achieve the optimum weapon system configuration by assessing fiber optic and hardwired systems for future applications.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):838-844. doi:10.1115/1.3239820.

This paper describes the control mode analysis procedure that is used to establish closed-loop control requirements for advanced aircraft propulsion systems. The procedure utilizes anticipated variations in engine component performance, engine deterioration, and control tolerances in a statistical analysis to establish corresponding variations in engine output performance and safety parameters. Potential closed-loop control configurations are evaluated by this process, compared, and the best configuration selected for implementation into control law and schedule designs. A byproduct of the analysis is the establishment of engine design and performance margins. The paper will describe typical engine variational models used in this process, the General Electric COMET program designed to automate the analysis, and typical mode study results based on a current augmented turbofan engine.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):845-850. doi:10.1115/1.3239821.

The design of multivariable control systems for modern applications is an important challenge to the control system engineer. Active control of metal machining operations, control of gas turbine operations, and chemical process control are current areas of interest. In precision machining operations where tolerances of a few microinches are required, in-process control with several servos will be required. These servos could control spindle axis location and tool position about several axes, thus forming a multi-input/multi-output system. Modern gas turbines, required to operate over more extended regimes, are provided with multiple controls, e.g., nozzle settings and fuel flows which must be implemented in some rational manner. In the chemical process industries, there are many examples of multivariable systems with several control variables and several desired or controlled outputs. One control approach considers a separate system for each of the controlled variables, so that a change in one input will produce an interaction effect that must be managed by another separate system. This approach is attractive and straightforward to implement, but current practice shows significant coupling effects. In order to reduce or eliminate interaction, a control algorithm, with strong integral compensation, for a sixth-order, two-input, two-output linear plant with dynamic coupling is proposed. Decoupling filters are not used. The primary goal is to realize a substantial reduction in the coupling effects when a step input is used for a single variable. A secondary goal is to achieve deadbeat response for the controlled variable to the step input. Moreover, these goals are to be attained in the presence of significant changes in the system parameters or in the presence of arbitrary external disturbances, i.e., robustness is required. The control strategy uses cascaded integral error compensation that permits conceptual division of the network so that two single-input/single-output systems result. Coupling effects are treated as arbitrary disturbances. Poles for each loop are placed on the real axis in the left-half plane. Numerical solutions to the system equations show that this approach produces a system that achieves effective decoupling and robustness simultaneously. It is significantly superior to Proportional-Integral-Derivative controllers which also are considered in detail.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):851-855. doi:10.1115/1.3239822.

The term “fiber optics” means the use of dielectric waveguides to transfer information. In aircraft systems with digital controls, fiber optics has advantages over wire systems because of its inherent immunity to electromagnetic noise (EMI) and electromagnetic pulses (EMP). It also offers a weight benefit when metallic conductors are replaced by optical fibers. To take full advantage of the benefits of optical waveguides, passive optical sensors are also being developed to eliminate the need for electrical power to the sensor. Fiber optics may also be used for controlling actuators on engine and airframe. In this application, the optical fibers, connectors, etc., will be subjected to high temperatures and vibrations. This paper discusses the use of fiber optics in aircraft propulsion systems, together with the optical sensors and optically controlled actuators being developed to take full advantage of the benefits which fiber optics offers. The requirements for sensors and actuators in advanced propulsion systems are identified. The benefits of using fiber optics in place of conventional wire systems are discussed as well as the environmental conditions under which the optical components must operate. Work being done under contract to NASA Lewis on optical and optically activated actuators sensors for propulsion control systems is presented.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):856-860. doi:10.1115/1.3239823.

On May 20, synthetic gas made by gasifying coal was burned in a commercial size gas turbine in Daggett, CA and the electricity produced from that coal flowed into the Southern California Edison system. This event marked the initial operational phase of the Cool Water Coal Gasification Program that expects to demonstrate that the integrated coal gasification combined cycle (IGCC) power plant can burn coal in an environmentally superior manner while also meeting a utility system’s operational and reliability requirements.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):861-869. doi:10.1115/1.3239824.

An ideal open gas turbine cycle with multiple-stage intercooled compression and multiple-stage reheat expansion theoretically approaches Carnot thermal efficiency. A proposed practicable process to utilize this cycle with indirect firing of coal as fuel, with an air heater in place of the boiler, a turbine inlet temperature of 1700°F (927°C) and top pressure of 788 psia (53.6 atm) gives promise of lowering power plant station heat rates (HHV) from 8970 Btu/kWh currently realized by the best scrubber-equipped coal fired steam plants to 7460, a reduction of 16.8% in fuel consumption and consequently the cost of flue gas scrubbers. In addition, a 316 MW plant delivers at rated output 130,000 gal per day water stripped from atmospheric air. Primarily because of an expensive air heater and regenerator the gas turbine plant is penalized by an estimated increase in initial cost from $ 1000/kW for a steam plant to $ 1433/kW. With coal priced at $ 3/million Btu, water selling at 2¢ /gal, money at 8% interest, inflation at 5% , and an 81% plant capacity factor, the payback period is 17 years.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):872-876. doi:10.1115/1.3239827.

Performance calculations for centrifugal compressors have been based on polytropic analysis for many years. The basic polytropic equation, in which head is found from gas pressure, temperature, and compressibility factor at the end points of compression, is applied by virtually all engineers involved with turbomachinery design, selection, or operation. The problems and errors associated with this simple calculation method when applied to nonideal gases have long been recognized. Schultz [1] proposed a correction factor to the head equation to compensate for the errors. This “polytropic head factor” correction is required by the ASME Power Test Code 10 [2] for the evaluation of compressor performance. Recently, the accuracy of even the corrected head equation has been questioned for the compression of gases to high pressures and an alternative calculation method has been proposed by Mallen and Saville [3]. Although differences were found between their method and the Schultz method of up to three percent, they did not show which method was more accurate. This paper evaluates the accuracy of these previous calculation methods and shows that both have errors for some compression calculations. In addition, a new polytropic calculation method is described and is shown to be substantially more accurate than the methods of both [1] and [3], thus allowing more precise evaluations of compressor performance.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):880-889. doi:10.1115/1.3239832.

The energy-balance approach to cycle analysis has inherent limitations. These arise from the fact that the first law of thermodynamics contains no distinction between heat and work and no provision for quantifying the “quality” of heat. Thus, while producing the correct final result, energy-balance analysis is incapable on its own of locating sources of losses. Identifying and quantifying those sources can be a useful design tool, especially in developing new, more complex systems. The second law of thermodynamics, applied in the form of entropy and availability balances for components and processes, can locate and quantify the irreversibilities which cause loss of work and efficiency. Perhaps one reason that such analysis has not gained widespread engineering use may be the additional complication of having to deal with the combustion irreversibility, which introduces an added dimension to the analysis. A method of cycle analysis, believed to circumvent this added difficulty for combustion cycles, is proposed and applied to complex combined cycles with intercooling and reheat. The fuel is treated as a source of heat, which supplies potentially available work to the cycle depending on the peak temperature constraint on work extraction. The availability is then traced as it cascades through the cycle, portions of it being wasted by the various components and processes, and the balance emerging as shaft work. Linkage with the traditional first-law efficiency is thus preserved, while establishing the location, cause, and magnitude of losses. Analysis and results for combined cycles with component irreversibilities are presented. The air-standard approximation with constant properties is used for simplicity. The turbine is treated as adiabatic since the cooling losses depend on the type of technology applied, particularly at higher temperatures. A model for quantifying those losses is presented in Part 2 of this paper.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):890-901. doi:10.1115/1.3239833.

Accurate heat rate functions for cyclically loaded coal-fired power plants are extremely difficult to obtain without the aid of plant computer systems which most older U.S. plants do not have. Special tests or theoretical heat balances generally are not the best means for developing these functions for reasons discussed. An easily applicable, but sufficient, regression-based method of determining the normal operating heat rate equation under these circumstances is presented. Along with the model used, actual results of application of the methodology to several plants are shown and discussed. Special emphasis is given to use for performance monitoring and incremental heat rate development.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):902-907. doi:10.1115/1.3239834.

This paper is concerned with the results of a theoretical investigation on combustion of traditional fuel and alcohol blends. An analytical procedure has been developed which examines three different hypotheses for introducing the alcohol: constant mass of primary fuel, constant total energy of fuel, and constant total mass of fuel. The procedure has been applied to combustion at constant volume varying over a wide range of air-fuel ratios, percentage of alcohol, and combustion temperature. The results obtained, of particular interest for reciprocating internal combustion engines, indicate that as far as energy and emissions are concerned, the effects of alcohol on combustion depend strongly on the hypothesis adopted for fueling the alcohol.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):908-913. doi:10.1115/1.3239835.

A series of tests was conducted on a Toyota, four-cylinder, spark ignition engine which was modified to run on either gasoline or natural gas. The aim of the experiment was to investigate the performance and combustion behavior of natural gas, with particular emphasis on its low burning velocity. A pressure transducer installed in the cylinder head was used to obtain pressure versus crank angle curves from which mass burn rates and burning velocities were calculated, using a heat release analysis program. Results indicate that the low laminar burning velocity of natural gas extends its ignition delay period (time to 1 percent burned) by up to 100 percent compared with gasoline. This contrasts with the remainder of the combustion period which is dominated by turbulence effects that produce very similar burning velocities for the two fuels.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):914-921. doi:10.1115/1.3239836.

The feasibility of dual-fuel operation with natural gas in a prechamber diesel engine was studied with special emphasis on fuel consumption and cylinder pressure development. The effects of air restriction, pilot diesel flow rate, and injection timing were also investigated. Near full load the fuel energy consumption rate was close to that of straight diesel operation though at part load (in the absence of air restriction) the fuel energy consumption rate was relatively high. In the absence of injection timing adjustment the maximum power output of dual-fuel operation was severely limited by the maximum cylinder pressure. Retarding the injection timing is effective in reducing the maximum cylinder pressure to a safe level. The analysis of apparent energy release indicates the differences in combustion mechanism between auto-ignition of diesel fuel in straight diesel operation and propagation of flame fronts in dual-fuel operation.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):922-930. doi:10.1115/1.3239837.

This paper describes the procedure developed by Cooper-Bessemer for large-bore gas engines to calculate the knock rating of gas fuel blends and to predict with accuracy the required engine build to use that fuel with optimum detonation margin. Engine prototype test work has included fuel sensitivity tests mapped as a function of compression ratio, fuel air ratio, ignition advance, combustion air temperature, and engine rating. Success in predicting production engine operation for a given application involving a particular fuel blend has been gratifying. The basic reference method blend selected was normal butane in methane. Details are included in the paper to illustrate the problems in making sensitivity correlations between small-bore fuel research engines and large-bore production engines.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):931-937. doi:10.1115/1.3239838.

This paper describes the design and testing of a high work capacity single-stage transonic turbine of aerodynamic duty tailored to the requirements of driving the high-pressure core of a low cost turbofan engine. Aerodynamic loading was high for this duty (ΔH/U 2 = 2.1) and a major objective in the design was the control of the resulting transonic flow to achieve good turbine performance. Practical and coolable blading was a design requirement. At the design point (pressure ratio = 4.48), a turbine total to total efficiency of 87.0 percent was measured—this being based on measured shaft power and a tip clearance of 1.4 percent of blade height. In addition, the turbine was comprehensively instrumented to allow measurement of aerofoil surface static pressures on both stator and rotor—the latter being expedited via a rotating scanivalve system. Downstream area traverses were also conducted. Analysis of these measurements indicates that the turbine operates at overall reaction levels lower than design but the rotor blade performs efficiently.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):938-944. doi:10.1115/1.3239839.

A special thin film-hardware material thermocouple (TC) and heat flux gauge concept for a reasonably high-temperature and high heat flux, flat-plate heat transfer experiment was fabricated and tested to gauge temperatures of 911 K. This unique concept was developed for minimal disturbance of boundary layer temperature and flow over the plates and minimal disturbance of heat flux through the plates. Comparison of special heat flux gauge Stanton number output at steady-state conditions with benchmark literature data was good and agreement was within a calculated uncertainty of the measurement system. Also, good agreement of special TC and standard TC outputs was obtained and the results are encouraging. Oxidation of thin film thermoelements was a primary failure mode after about 5 hr of operation.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):945-952. doi:10.1115/1.3239840.

Fully three-dimensional periodic flows through a turbine stage of stator and rotor are studied numerically by solving time-dependent three-dimensional Euler equations with the finite-volume method. The phase relation of stator and rotor flows and the related blade-row interaction are accounted for in the time-space domain. The established method of numerical calculation makes a practical contribution to predict actual turbine flows through a turbine stage of stator and rotor which have an arbitrary number of blades.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):953-960. doi:10.1115/1.3239841.

Local heat-transfer coefficients were experimentally mapped along the midchord of a five-times-size turbine blade airfoil in a static cascade operated at room temperature over a range of Reynolds numbers. The test surface consisted of a composite of commercially available materials: a mylar sheet with a layer of cholesteric liquid crystals, which change color with temperature, and a heater sheet made of a carbon-impregnated paper, which produces uniform heat flux. After the initial selection and calibration of the composite sheet, accurate, quantitative, and continuous heat-transfer coefficients were mapped over the airfoil surface. The local heat-transfer coefficients are presented for Reynolds numbers from 2.8×105 to 7.6×105 . Comparisons are made with analytical values of heat-transfer coefficients obtained from the STAN5 boundary layer code. Also, a leading-edge separation bubble was revealed by thermal and flow visualization.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):961-968. doi:10.1115/1.3239842.

The loss mechanisms and the behavior of secondary flows downstream of a large scale, linear turbine cascade have been investigated experimentally. A five-blade replica of the cascade used by Langston et al. at United Technologies Research Center was used for the present tests. Detailed flow measurements, using five-hole and three-hole probes, were made at four different planes, one just upstream of the trailing edge and the rest downstream. The secondary flow field at each measurement plane was found to be dominated by a single large passage vortex, which decayed in strength because of the mixing occurring in the flow. More than one-third of the losses were found to occur downstream of the trailing edge. This rise in total pressure loss in the present tests was almost entirely explained by a corresponding dissipation of the secondary kinetic energy of the flow. A mixing analysis of the flow was done to predict the additional losses due to “mixing” until the flow became completely uniform.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):969-975. doi:10.1115/1.3239843.

The overall performance of two geometrically similar linear turbine cascades is calculated using an elliptic flow program. The increase in the mass-averaged total pressure loss is calculated within and downstream of the cascades and the results show good agreement with the measured values. The buildup and decay of the secondary kinetic energy are also shown; measurements are available for one of the cascades near and downstream of the trailing edge and these are in close agreement with the calculated values. Details of the flow development are also compared with measurements. Calculated velocity vectors near the endwall show the overturning revealed by surface flow visualization and similarly near the suction surface the strong spanwise flow is well calculated. Calculated contours of total pressure loss in cross-sectional planes confirm the important interaction of the passage vortex with the profile boundary layer at midspan. Regions of high loss near midspan are calculated downstream of both cascades; this three-dimensional flow development is followed in the calculations.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):976-982. doi:10.1115/1.3239844.

The effects of sinusoidal flow pulsations on the heat transfer from a cylinder to a crossflow at Re = 50,000 were investigated. A range of different pulsation amplitudes of up to 25% and frequencies both above and below the natural shedding frequency were used. The pulsating flow was clean and well organized. It had greater than 95% of the power at the fundamental frequency with a low turbulence level (less than 0.5%). The time-averaged local heat transfer was experimentally measured for a constant-temperature surface-boundary condition using a small heat flux gage in the cylinder wall. Distributions were obtained by rotating the cylinder through 180 deg. The experiments showed no significant increase of heat transfer due to the flow pulsation in either the wake or attached boundary layer region. Small local increases were found near the separation point.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):983-990. doi:10.1115/1.3239845.

Tip endwall contouring is one of the most effective methods to improve the performance of low aspect ratio turbine vanes [1]. In view of the wide variety of geometric parameters, it appears that only the physical understanding of the three-dimensional flow field will allow us to evaluate the probable benefits of a particular endwall contouring. The paper describes the experimental investigation of the three-dimensional flow through a low-speed, low aspect ratio, high-turning annular turbine nozzle guide vane with meridional tip endwall contouring. The full impact of the effects of tip contouring is evaluated by comparison with the results of a previous study in an annular turbine nozzle guide vane of the same blade and cascade geometry with cylindrical endwalls [12]. In parallel, the present experimental study provides a fully three-dimensional test case for comparison with advanced theoretical calculation methods [15]. The flow is explored by means of double-head, four-hole pressure probes in five axial planes from far upstream to downstream of the blade row. The results are presented in the form of contour plots and spanwise pitch-averaged distributions.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):991-997. doi:10.1115/1.3239846.

This paper deals with an experimental investigation of heat transfer across the suction side of a high-pressure, film-cooled gas turbine blade and with an attempt to numerically predict this quantity both with and without film cooling. The measurements were performed in the VKI isentropic compression tube facility under well-simulated gas turbine conditions. Data measured in a stationary frame, with and without film cooling, are presented. The predictions of convective heat transfer, including streamwise curvature effects, are compared with the measurements. A new approach to determine the augmented mixing lengths near the ejection holes on a highly convex wall is discussed and numerical results agree well with experimentally determined heat transfer coefficients in the presence of film cooling.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):998-1006. doi:10.1115/1.3239847.

The unsteady effects of shock waves and wakes shed by the nozzle guide vane row on the flow over a downstream turbine rotor have been simulated in a transient cascade tunnel. At conditions representative of engine flow, both wakes and shock waves are shown to cause transient turbulent patches to develop in an otherwise laminar (suction-surface) boundary layer. The simulation technique employed, coupled with very high-frequency heat transfer and pressure measurements, and flow visualization, allowed the transition initiated by isolated wakes and shock waves to be studied in detail. On the profile tested, the comparatively weak shock waves considered do not produce significant effects by direct shock-boundary layer interaction. Instead, the shock initiates a leading edge separation, which subsequently collapses, leaving a turbulent patch that is convected downstream. Effects of combined wake- and shock wave-passing at high frequency are also reported.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):1007-1015. doi:10.1115/1.3239804.

Experimental results are presented to document hydrodynamic and thermal development of flat-plate boundary layers undergoing natural transition. Local heat transfer coefficients, skin friction coefficients, and profiles of velocity, temperature, and Reynolds normal and shear stresses are presented. A case with no transition and transitional cases with 0.68 percent and 2.0 percent free-stream disturbance intensities were investigated. The locations of transition are consistent with earlier data. A late-laminar state with significant levels of turbulence is documented. In late-transitional and early-turbulent flows, turbulent Prandtl number and conduction layer thickness values exceed, and the Reynolds analogy factor is less than, values previously measured in fully turbulent flows.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):1016-1021. doi:10.1115/1.3239805.

This paper describes an experimental heat transfer investigation around the leading edge of a high-pressure film-cooled gas turbine rotor blade. The measurements were performed in the VKI isentropic compression tube facility using platinum thin film gauges painted on a blade made of machinable glass ceramic. Free-stream to wall temperature ratio, Reynolds, and Mach numbers were selected from actual aeroengines conditions. Heat transfer data obtained without and with film cooling in a stationary frame are presented. The effects of coolant to free-stream mass weight ratio and temperature ratio were successively investigated. Heat transfer modifications due to incidence angle variations were interpreted with the aid of inviscid flow calculation methods.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):1022-1030. doi:10.1115/1.3239806.

The effect of the interaction of the wake from a nozzle guide vane with the rotor may be simulated in part by means of a stationary rotor and a moving wake system. This technique is applied to a transonic rotor blade cascade, and the unsteady measurements of surface pressure and heat transfer rate are compared with baseline data obtained without the wake interaction. The wake-rotor interaction results in a change in inlet incidence angle and this effect is also examined in the steady-state case. It is found that the shock waves from the moving wake system have a major effect on the instantaneous heat transfer rates.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):1031-1038. doi:10.1115/1.3239807.

Experimental measurements of the flow field in a low-speed, large-scale, annular cascade of highly loaded turbine rotor blades are presented. The blade has a turning angle of 128.5 deg, an aspect ratio of 0.88, and a Zweifel coefficient of 1. Detailed cascade tests consisted of inlet and exit flow parameter traverses, blade passage pressure distributions, and flow visualization. The results are presented in the form of contour plots and pitch-averaged radial distributions of losses and flow angles. The measurements are compared with the results obtained for the same blade section tested in a planar cascade. Distribution of the losses and flow angles revealed the presence of two large vortices that occupied a major portion of the trailing edge plane. A large high-loss core was visible in the center of the blade passage and coincided with regions of maximum flow underturning. The measured cascade secondary losses compared well with existing correlations.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 1985;107(4):1039-1046. doi:10.1115/1.3239808.

Measurements of the unsteady flow field near and within a turbine rotor were made by means of a Laser-2-Focus velocimeter. The testing was performed in a single-stage cold-air turbine at part-load and near-design conditions. Random unsteadiness and flow angle results indicate that the secondary vortices of the stator break down after being cut and deformed by the rotor blades. A quantitative comparison shows that some of the energy contained in these secondary vortices is thereby converted into turbulence energy in the front part of the rotor. An attempt is made to explain this turbulence energy production as caused by the vortex breakdown.

Commentary by Dr. Valentin Fuster

DISCUSSIONS

Commentary by Dr. Valentin Fuster
Commentary by Dr. Valentin Fuster

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