This paper focuses on the influence of shaft labyrinth seal flow on full stage performance. Experimental data are studied, expected design conditions and experimental results are compared and discussed and a losses breakdown for the design procedure is presented. The experimental investigation was performed in VZLU’s air test turbine which is a part of a closed-loop system equipped with a radial compressor. The test turbine configuration simulated the real drum-stage geometry of an axial steam turbine. The geometry of the turbine represents a typical mid-pressure stage of a steam turbine. The configuration of the test rig was adapted in order to easily change the shaft labyrinth seal geometry. The study covered a wide range of seal clearances from very small to extremely large clearances, reaching a maximum relative mass flow approximately 10% of the stator blade flow. Different types of seal feed were also tested to compare internal feed (the flow obtained from the stator flow by the hub-gap just in front of the stator) and external feed realized by additional piping with external regulation. Three stage reactions were tested in this work — Low Reaction, Mid Reaction and Full Reaction. The stator of the stages was the same in all cases, thus the reaction was changed by implementing three different rotor geometries. The influence of the labyrinth seal clearance was investigated by overall performance measurement and by detailed investigation of the flow field. The turbine stage was loaded by a hydraulic dynamometer used for regulating the rotational speed and a flange torquemeter was used to determine the stage efficiency. The total mass flow was measured using an orifice plate. Each seal geometry configuration was calibrated to compute the seal mass flow. The turbine stage and seal were equipped with a number of static pressure taps, and miniature pressure probes were used for measuring the flow field parameters in detail. The discussion of the results is divided into two areas. Firstly, the influence of the degree of reaction on axial steam turbine stage performance in the configuration without the seal flow is presented. Then, a combination of various degrees of reaction is studied as a function of mass flow through the shaft labyrinth seal. The measured data are evaluated by a breakdown of loss sources. The decomposition of the total loss into row losses, leakage losses and mixing losses is highly advantageous. This total loss analysis is carried out for all three stages and both off-design performance and ratios of the shaft seal flow to nozzle blade flow are measured. The post-processing of measured data through this loss breakdown and the comparison with the design is used to validate the design process.

This paper presents the development and validation of a meanline model by means of numerical and experimental methods, to determine it’s feasibility as an optimisation tool for turbocharger matching. Using a parametric turbine model, numerical experiments were conducted accounting for variations of several key turbine design parameters and a wide operating range. The resulting dataset was used to test the accuracy of the meanline model when calibrated to a baseline design and thus evaluate it’s ability of extrapolating to different designs. The loss models were examined in more detail, and a set of loss models which provided the most accurate results is presented. The meanline model was further validated experimentally using dynamometer test results of 6 turbine designs from the same parametric turbine model. The result showed that for design point and high power operation, an error of less than 3.1% and 2.0% was achieved for efficiency and mass flow parameter respectively. This led to the conclusion that the model would be sufficiently accurate to represent design changes relevant to turbocharger matching.

A gas turbine engine has supported the U.S. Military Academy’s mechanical engineering program for nearly three decades. Recent, substantial enhancements to the engine, controls, and data acquisition systems greatly increased the student experience by leveraging its broad capabilities beyond the original laboratory learning objectives. In this way, the laboratory served as a learning platform for more than just instruction on gas turbine fundamentals and the Brayton cycle. The engine is a refurbished auxiliary power unit from Pratt & Whitney Aeropower, installed in the Embrauer 120 and similar to a unit installed on a U.S. Army helicopter. Whereas the original laboratory experience permitted students to test the engine at three different loads applied by a water brake dynamometer, the revised experience allowed for a broader range of test conditions. The original laboratory included single point measurements of three temperatures and two pressures, along with the fuel flow rate, dynamometer torque, and engine speed. The revised laboratory allowed the user to vary bleed air and engine loads across an operational envelope at a user-specified acquisition rate. The improved data acquisition system used LabVIEW™ and included multiple state sensors for pressure, temperature, fuel flow, bleed air, and dynamometer performance, thereby enabling a more complete analysis by accounting for the energy transported by bleed airflow and absorbed by the water brake. Students then quantified the uncertainty in their measurements and analysis. The new emphasis on uncertainty quantification, part of a program-level initiative, challenged students’ notion of “substitute and solve” while also familiarizing them with large, experimental data sets. The re-envisioned laboratory raised the students’ level in the cognitive domain and served as their premier engine experience. Rather than merely observing engine adjustments across a small range of conditions, students designed their own laboratory experience. With the updated approach, students viewed a graphic of the turbine’s laboratory operating range and chose the key variables of interest — selecting data points within the laboratory operating range — and then justified their selections. The enhanced experience added analysis of flow exergy and exergetic efficiency. The exercise also challenged students to hypothesize why actual turbine performance was less than predicted and determine sources of error and uncertainty. Moreover, the new laboratory offers opportunities to expand the turbine engine’s utility from supporting a single thermal-fluids course to a multidisciplinary learning platform. Concluding remarks address concepts for augmenting course instruction in other courses within the curriculum, including heat transfer, mechanical vibrations, and dynamic modeling and controls.

The design of a low reaction turbine blade profiles was carried out to improve the steam flow efficiency. The blade profiles geometry design for both the stationary and moving blades and reprofiling of them are done using Vista ATBlade, according to the aerodynamic analysis results from the cascade analysis code MISES. The original stator profile is aft-loaded, and the new one present in this paper is highly-aft-loaded (HAL) to depress the development of secondary flow further, while maintaining even lower profile loss and wider incidence angle tolerance. The newly designed moving blade is more robust compared with the original one, thus it has larger aspect ratio under the same blade section average stress level, and with better incidence tolerant capability as well. The planar cascade air tests were first carried out to verify the stator profile loss improvement, with a decrease of energy loss coefficient of almost 0.8% obtained under the Reynolds number of about 1e6. Then the annular cascade air tests with fully 360 degrees stator blades installed were conducted to validate the reduction of endwall loss and the profile loss as well, and to measure the mass flow capability (real mass flow/ideal mass flow). Finally, two three-stage tests for the original blades and the new one were developed to verify the improvement under real multi-stages flow conditions. All the stages for both tests are designed with the hub reaction of about 15%, without interstage swirl, in the design condition. The flow probes at upstream of first stage stator and downstream of last stage moving blade, the hydraulic dynamometer and the flowmeter are used to test the overall efficiency. Three traverse planes are located at the upstream, middle and downstream of the second stage to measure the flow properties using five hole pneumatic probes. The test results showed a increase of overall efficiency of about 1.5%. The CFD simulations showed very good agreement of mass flow capability with the tests, for both the stator annular and multi-stage tests. The application of the newly designed blade profiles in SanHe subcritical reheat 300MW steam turbine (16.7MPa/537°C/537°C) retrofit gives the final proof of the efficiency improvement. The measured efficiency showed remarkable performance, with an increase of efficiency of 1.5%–2.2% for both the HP and IP cylinder.

In conjunction with the SunShot program to develop a supercritical CO 2 turbine expander, an air dynamometer was developed to absorb the power of the SCO 2 turbine during development testing. The dynamometer is unique in that it is directly mounted to the shaft of the turbine expander replacing the associated compressor coupling during the test campaign. It is able to replicate both the load of the compressor as well as the rotordynamic qualities of the compressor coupling all without the need for additional bearings, casing, or larger skid. The dynamometer consists of a single-stage centrifugal air compressor ingesting ambient air. To maximize turn-down and minimize weight, the impeller diameter and flow coefficient were optimized along with the design of the loop throttle valves. The detailed aerodynamic, rotordynamic, and mechanical design for the dynamometer is presented. The aerodynamic methodology to minimize unsteady forces transmitted to the turbine will also be reviewed.

Increasing engine cycle pressure ratios and temperatures, and assuring improved component reliability and life, have led to complex blade cooling systems which, in part, compromise the aerodynamic design of today’s high pressure (HP) turbines in aero engines. The National Gas Turbine Establishment is currently constructing a new research module, and associated plant and services, to investigate the performance of cooled HP turbines. The unit will satisfy the requirement to measure the power developed by high temperature turbines when operating in an engine environment with all the cooling flows properly represented and monitored accurately. Research on the thermal design of rotor blade cooling systems will also be undertaken at realistic temperatures, pressures and turbulence level, and with correct rotational effects. The module is instrumented extensively to determine mainstream, coolant and leakage flow characteristics, and the power will be absorbed by a dynamometer connected through a torquemeter. Data on the rotor blade metal temperature, surface pressure, and coolant pressure and temperature will be transmitted via telemetry and rotating scanivalve. Data acquisition and control for the high temperature research module will be by the use of a mini computer-logger system. This paper outlines the concept and design considerations of the high temperature turbine research module. The instrumentation of the module is also presented and discussed.

Testing of turbo-shaft engines requires power absorption devices that must be capable of stable operation at both variable powers and speeds (idle, take-off, cruise). A dynamometer transfers the engine shaft power to the incoming fluid and transforms it into thermal energy. An air-dynamometer has many advantages and is preferred over water brakes or electric generators, but must be made to operate stably over a wide range of speeds and powers.

This paper describes the development of a ceramic turbocharger at Ford Motor Company, Engineering and Research Staff. The potential advantages of reduced turbocharger and vehicle response time and manufacturing cost savings have made the ceramic turbocharger rotor a subject of interest to automotive, ceramic, and turbocharger manufacturers. An Ishikawajima-Harima Heavy Industries Co., Ltd. RH06 turbocharger was modified to incorporate a slip cast reaction bonded silicon nitride rotor to evaluate the effect of reduced rotary inertia on vehicle performance. The results of vacuum spin pit testing, turbocharger test cell testing, 2.3L spark ignition dynamometer tests, and preliminary vehicular tests, are reported.

In advance of initial dynamometer testing of the AGT 100 engine, all prime components and subsystems were bench/rig tested. Included were compressor, combustor, turbines, regenerator, ceramic components, and electronic control system. Results are briefly reviewed. Initial engine buildup was completed and rolled-out for test cell installation in July 1982. Shakedown testing included motoring and sequential firing of the combustor’s three fuel nozzles.

Turbocharger “lag” or poor response to engine load changes can be improved by reducing the rotating inertia of the turbocharger turbine, compressor and shaft system. Recently designed, second generation turbochargers all have small diameter, light weight rotating assemblies in an effort to minimize inertia and improve response. An automotive turbocharger with an axial flow turbine rather than a conventional radial inflow turbine is presented here as an alternative method of reducing inertia. The rotating inertia of the axial flow turbine and a centrifugal compressor is about one half that of the same compressor combined with a radial inflow turbine. In steady-state engine dynamometer tests, the same wide-open throttle performance was obtained with both turbochargers. Engine dynamometer transient tests showed that the turbocharger with the axial flow turbine attained full boost 25–40% faster than did the turbocharger with the radial inflow turbine.

Pratt & Whitney Canada produces a wide range of aircraft engines and this has led to a concerted effort to standardize and streamline its production engine test facilities. P&WC produce two very different series of turboprop engines, the PW100 with a conventional intake and exhaust arrangement and the PT6 with its reverse flow arrangement. A dynamometer test cell capable of testing both these engine series has been designed and built at Longueuil and is now in operation. The changeover from one model to the other can be carried out by an operator in less than two hours and requires no special tooling or manpower. This paper discusses the solutions developed to overcome the inherent problems of intake and exhaust arrangement, engine mounting, slave equipment requirements etc. generated by testing two very different families of engines in the same test cell coupled with the need to incorporate the efficiency and ease of operation required of a production facility.

This paper presents results of a program to investigate the magnitude, origin and parametric variations of destabilizing forces which arise in high power turbines due to blade-tip leakage effects. Five different unshrouded turbine configurations and one configuration shrouded with a labyrinth seal were tested with static offsets of the turbine shaft. The forces along and perpendicular to the offset were measured directly with a dynamometer, and were also inferred from velocity triangles and pressure distributions obtained from detailed flow surveys. These two routes yielded values in fair agreement in all cases. For unshrouded turbines, the cross-forces are seen to originate mainly (∼2/3) from the classical Alford mechanism (nonuniform work extraction due to varying blade efficiency with tip gap) and about 1/3 from a slightly skewed hub pressure pattern. The direct forces arise mainly (3/4) from this pressure pattern, with the rest due to a slight skewness of the Alford mechanism. The pressure nonuniformity (lower pressures near the widest gap) is seen to arise from a large-scale redistribution of the flow as it approaches the eccentric turbine. The cross-forces are found to increase substantially when the gap is reduced from 3.0% to 1.9% of blade height, probably due to viscous blade-tip effects. The forces also increase when the hub gap between stator and rotor decreases. The force coefficient decreases with operating flow coefficient. In the case of the shrouded turbine, most of the forces arise from nonuniform seal pressures. This includes about 80% of the transverse forces. The rest appears to come from uneven work extraction (Alford mechanism). Their level is about 50% higher than in the unshrouded cases.

Performance evaluations of rocket engine turbopump drive turbines are difficult to obtain from turbopump or engine firings due to measurement limitations and operating point restrictions. The Marshall Space Flight Center (MSFC) Turbine Test Equipment (TTE) was developed to provide an accurate, economical method of measuring the performance of full-scale turbopump gas turbines. By expanding air at pressures as high as 435 psia (3.0 MPa) to atmospheric conditions, the TTE provides metered air at nominal conditions of 100 psia (0.69 MPa), 550 °R (350 °K), and 15 lbm/sec (6.8 kg/sec) with run times of 100 seconds or greater. A 600 hp (448 kW) direct current dynamometer and gearbox provide turbine power absorption for speeds up to 14,000 rpm. This paper describes the MSFC TTE and its performance including the performance envelope, turbine inlet flow quality, and measurement uncertainty.

Combustor hardware employing catalytic combustion technology has been developed for a 1.5 MW gas turbine. This system, combined with state of the art catalyst technology, was used to demonstrate ultra-low emissions on the engine. The demonstrator combustor utilizes a two stage lean premix preburner system to obtain the required catalyst inlet temperatures and low NOx over the operating load range. The performance of the preburner system was characterized during engine tests by measuring temperature rise and emissions just downstream of the preburner. A fuel schedule for the primary and secondary stages was selected to give NOx emissions below 2 ppmv at the engine exhaust. Overall engine performance was measured over the full load range. Emissions of NOx < 3 ppmv and CO and UHC < 5 ppmv were obtained at 72% to 100% load. Combustor dynamics were shown to be less than 0.3 psi(rms). This combustor operated for 1000 hours on a dynamometer test facility and showed low emissions performance over this period.

A study of gas turbine engines is an important component of an integrated thermodynamics and fluid mechanics two-course sequence at the United States Military Academy (USMA). Owing to the ubiquity of gas turbines in military use, graduating cadets will encounter a variety of these engines throughout their military careers. Especially for this unique population, it is important for engineering students to be familiar with the operation and applications of gas turbines. Experimental analysis of a functional auxiliary power unit (APU) from an Army utility helicopter has been a key component of this block of instruction for several decades. As with all laboratory equipment, the APU has experienced intermittent maintenance issues, which occasionally render it unusable for the gas turbine laboratory in the course. Because of this, a very basic virtual laboratory was implemented which integrated video of the physical laboratory with key parameters and behind-the-screen data collection for use in engine analysis. A revitalized version of both the physical and virtual gas turbine laboratory experiences offered in the thermal-fluids course will include substantial improvements over the existing setup. The physical laboratory, which is centered on a refurbished APU from a medium-sized commercial aircraft, will continue to incorporate measurements of temperature and pressure throughout the combustion process, as well as fuel flow rate. In an improvement over the original laboratory setup, an orifice plate will be used to measure the flow rate of bleed air exiting the turbine, which had not previously been open during engine testing. Additionally, the air flow through the anti-surge valve was not metered in the original version of the physical laboratory. However, the anti-surge air flow can account for nearly 25% of the total air flow, and performance calculations in the physical laboratory will now account for this loss. The turbine output shaft will run a water-brake dynamometer. All instrumentation will be converted to digital signals and projected on a large screen outside the test area through a LabVIEW front panel. The virtual laboratory will include the same metering options as the operational APU. In addition, the virtual laboratory will include the option to alter engine operating parameters, such as inlet temperature and pressure or exhaust temperatures, and students may conduct broad parameter sweeps across ranges of possible inputs or desired outputs. These improvements will enable students to gain a deeper understanding of gas turbine operation and capabilities in practical applications. The improved laboratory will be implemented in Spring, 2014.

The ultimate goal of an advanced turbocharger development is to have a superior aerodynamic performance while having the turbocharger survive various real world customer applications. Due to the uncertainty of customer usage and driving pattern, the fatigue life prediction is considered one of the most ambiguous analyses in the entire design and analyses processes of the turbocharger. The turbocharger system may have various resonant frequencies, which may be within the range of turbocharger operation for automotive applications. A turbocharger may operate with excessive stresses when running near resonant frequencies. The turbocharger may experience fatigue failures if the accumulative cycles of the turbocharger running across the resonant frequencies exceeds a certain limit. In this study, the authors propose an alternative approach to mitigate this kind of fatigue issues: i.e. engine system approach to improve turbocharger fatigue life via avoiding operating the turbocharger near resonant speeds for extended period of time. A preliminary numerical study was made and presented in this paper to assess the feasibility of such an engine system approach, which is followed by an engine dynamometer test for engine performance sensitivity evaluation when the turbocharger operation condition was adjusted to improve the high cycle fatigue life. The study shows that for a modern diesel engine equipped with electrically controlled variable geometry turbine and EGR for emission control, through the engine calibration and control upgrade, turbocharger operation speed can be altered to stay away from certain critical speeds if necessary. The combined 1D and 3D numerical simulation shows the bandwidth of the turbine “risk zone” near one of the resonant speeds and the potential impact on engine performances if the turbocharger speed has to be shifted out of the “risk zone.”

Heavy EGR required on diesel engines for future emission regulation compliance has posed a big challenge to conventional turbocharger technology for high efficiency and wide operation range. This study, as part of the U.S. Department of Energy sponsored research program, is focused on advanced turbocharger technologies that can improve turbocharger efficiency on customer driving cycles while extending the operation range significantly, compared to a production turbocharger. The production turbocharger for a medium-duty truck application was selected as a donor turbo. Design optimizations were focused on the compressor impeller and turbine wheel. On the compressor side, advanced impeller design with arbitrary surface can improve the efficiency and surge margin at low end while extending the flow capacity, while a so-called active casing treatment can provide additional operation range extension without compromising compressor efficiency. On the turbine side, mixed flow turbine technology was revisited with renewed interest due to its performance characteristics, i.e. high efficiency at low-speed ratio, relative to the base conventional radial flow turbine, which is relevant to heavy EGR operation for future diesel applications. The engine dynamometer test shows that the advanced turbocharger technology enables over 3% BSFC improvement at part-load as well as full-load condition, in addition to an increase in rated power. The performance improvement demonstrated on engine dynamometer seems to be more than what would typically be translated from the turbocharger flow bench data, indicating that mixed flow turbine may provide additional performance benefits under pulsed exhaust flow on an internal combustion engine and in the low-speed ratio areas that are typically not covered by steady state flow bench tests.

This paper presents the development of a magnetoelastic torque-meter for use on the Landing Craft Air Cushion (LCAC) hovercraft’s high speed gas turbine engines. As the gas turbine can produce in excess of the nominal torque limit of the right angle gearbox, torque limiting is required. Limiting torque based on torque tables (as is done currently) penalizes the overall performance of the majority of the fleet due to a significant variance in engine output horsepower. This sub-optimal operation of the craft could be overcome by measuring actual torque produced by each engine with a torque-meter; however conventional torque-meter designs were deemed impractical to retrofit due to the design requirements, and issues with integration and reliability / maintainability. An in-depth search by the US Navy identified a torque-meter concept utilizing magnetoelastic polarized band technology. The aim of this paper is to: (i) outline the general architecture of the system, (ii) highlight the performance of the torque-meter developed for the LCAC, (iii) describe the efforts that helped to mature this technology to Technology Readiness Level 8 and to transition it from motorsport applications to use on the LCAC, and (iv) summarize the initial results of the torque-meter system validation obtained through dynamometer tests on the engine and craft tests on the LCAC.

This paper summarizes experimental results of an aerodynamic performance study carried out on two full stage turbine test rigs. The stage under investigation was designed as a gas generator turbine for a small jet engine produced by PBS (the TJ100 engine with thrust of 1100 N and turbine tip diameter of 141 mm). The investigation was carried out alternatively on two full-stage test rigs (in-scale and scaled-up) integrated into a cool flow closed-loop wind tunnel located at VZLU. Firstly, the in-scale testing, focusing on an overall stage performance measured by means of a hydraulic dynamometer was arranged. Furthermore, some time-averaged flow field parameters in terms of total pressure, velocity and angles were acquired along the channel height upstream and downstream of the stage. The flow path authenticity and construction simplicity were strictly followed during the rig design phase and therefore original parts of the engine were mostly used. Then, the verification of results was performed with the stage scaled-up by factor 2.27. The overall stage performance was measured and compared with results of the in-scale measurement. Moreover, detail unsteady flow field measurement at the rotor exit was performed. Time-resolved data were analysed in order to study the influences of the stage load on the stage performance.

Replacing a Brayton cycle near constant-pressure combustor with a pulsed detonation combustor (PDC) may take advantage of potential performance improvements from low-entropy, pressure-gain heat addition. In this paper, the radial turbine of a Garrett automotive turbocharger is coupled to a hydrogen fueled PDC. Unsteady turbine power is obtained with a conventional dynamometer technique. Sampling frequencies greater than 10 kHz resolve rapid flowfield transients of confined detonations which occur in less than a millisecond and include peak gas pressures exceeding 4 MPa and peak gas temperatures greater than 2,400 K. Results include 6 ms time histories of turbine inlet and exit temperature, pressure, mass flow, and enthalpy during blowdown of a PDC. The unsteady inlet flowfield included momentary reverse flow, which was not observed at the turbine exit. Full pulsed detonation cycle time histories of turbine power, rotor speed, rotational energy and net shaft torque are included to describe the turbine response to detonations. Rotor speed is periodic and net shaft torque oscillates in response to a detonation. Results are shown for fill fractions ranging from 0.5 to 1.0 with a 0.5 purge fraction. PDC operating frequencies in this study range from 10 Hz to 25 Hz.