Pressure Gain Combustion (PGC) is considered a possible solution to increase gas turbine cycle efficiency, due to the lower entropy generation in the combustion process. However, the highly unsteady flow produced by PGC makes it more difficult to extract work from its exhaust gas. Any outlet restriction downstream of PGC, such as turbine blades, affects its flow field and may cause additional thermodynamic losses. The unsteadiness in the form of pressure, temperature and velocity vector fluctuations has a negative impact on the operation of conventional turbines. Therefore, evaluating early turbine design parameters for such applications is of great interest. Additionally, experimental measurements and data acquisition present researchers with challenges that have to do mostly with the high temperature exhaust of PGC and the high frequency of its operation. Numerical simulations can provide important insights into PGC exhaust flow and its interaction with turbine blades. In this paper, a Rotating Detonation Combustor (RDC) and a row of nozzle guide vanes have been modeled based on the data from literature and an available experimental setup at TU Berlin. Five guide vane configurations with different geometrical parameters have been modeled. URANS simulations were done for all guide vane arrangements to investigate the effect of solidity and blade type representing different outlet restrictions on the RDC exhaust flow. Total pressure loss and velocity fluctuation were computed upstream and downstream of the vanes. The results analzed the connection between total pressure loss and the vanes solidity and thickness to chord ratio. It is observed that more than 57% of the upstream velocity angle fluctuation amplitude was damped by the vanes. Furthermore, the area reduction was found to be the significant driving factor for damping the velocity angle fluctuations, whether in the form of solidity or thickness on chord ratio increment. A further study of the flow field details revealed that the vane passages act as convergent divergent nozzles in the unsteady flow field and no compression wave exists upstream. This RDC exhaust flow investigation is an important primary step from a turbomachinery standpoint, which provided details of blade behavior in such an unsteady flow field.

During the static testing of a single spool turbojet engine (400 N Titan Engine) running on Jet A-1 as fuel, the starting sequence was unsuccessful, however, upon substituting the fuel with Diesel the engine started successfully. Till that point, the engine operation/performance was normal, and a series of tests were conducted for a total time of 5 hours. These 5 hours of testing spanned over five months during which ambient temperature varied from 5 °C to 37 °C, and relative humidity varied from 40% to 90%. The starting sequence was unsuccessful at the maximum ambient temperature (37 °C) and moderate relative humidity condition (50%). During the test, various parameters including pressure, temperature, rotor speed and mass flow rate were measured and recorded using a computer based data acquisition. For the failed start-up case, during the starting sequence, a sudden decrease in rotor speed was observed as the rotor reached ∼80% of the calibration value followed by flameout in combustion chamber. This sudden drop in rotor speed occurred due to mild-surge in compressor which continued up to flameout. The cause of mild-surge is attributed to the momentary blockage at the exit of the compressor caused by the combustion process. These pressure fluctuations propagated downstream leading to flame out. The combination of high ambient temperature (35 °C) and moderate humidity (∼50%) resulted in higher combustion chamber temperature which resulted in the momentary blockage at compressor exit during the transient operation. The substitution of Diesel as fuel resulted in lower combustion chamber temperature, and resolved the issue. It is important to note that, the mild-surge observed in this case is caused by downstream condition, and would occur only at some special situations (inlet conditions). Hence, it may be concluded, that the compressor surge in a jet engine may be caused by downstream components (under some special conditions), and design modifications to these downstream components may resolve the issue in those cases.

With the well-known Transonic Compressor Darmstadt (TCD) in operation since 1994, profound knowledge in designing and operating a sophisticated test-rig is available at the Institute of Gas Turbines and Aerospace Propulsion of TU Darmstadt. During this period, TCD has been subject to a vast number of redesigns within different measurement campaigns (see [1], [2], [3], [4], [5], [6], [7], [8]). To expand the capabilities and ensure a sustainable process of compressor research, a new test facility was designed and built by the institute. The new test rig Transonic Compressor Darmstadt 2 (TCD2) features increased power for higher pressure ratios and higher mass-flow, a state of the art control system, increased flexibility towards different compressor geometries and modern data acquisition hardware and software. Following the successful commissioning of the test-rig in March 2018, a first measurement campaign has been conducted. Early test results regarding aerodynamic performance and aeroelastic effects of the test compressor are presented together with a detailed overview of test-rig infrastructure and control systems as well as the test compressor and the measurement hardware.

The present study is an experimental investigation of the nature of acoustically induced flashback in a lab-scale dump combustor. The control parameters varied include the inlet Reynolds number ( Re ) and the inlet turbulence intensity. The primary bifurcation plots of the combustor from stable to the unstable condition are seen to be significantly altered by the inlet turbulence intensity, with the latter delaying the onset of combustion instability to higher Re . The analysis of multivariate high-speed data acquisition and processing (viz. unsteady pressure, flame imaging and velocity field by means of PIV) reveals the role of low-frequency high amplitude acoustics in modulating the flame. It is seen that high amplitude oscillations are sustained by two mechanisms 1. Modulation of the flame by coherent structures shedding at the step and 2. The bulk flame motion in-and-out at the edge of the step. It is seen that flow reversal at sufficiently low frequencies provide enough duration for the hot products to ignite fresh reactants upstream of the duct, which in-turn reinforces the coherent unsteadiness in the system, thereby increasing the propensity of the mixture to be ignited more upstream with every cycle. This ultimately leads to the flame flashing back till the point of premixing. This work thus addresses and reforms the occurrence of flashback being an example of loss of static stability, whereby the overriding presence of dynamic combustion instability results in a flashback to behave in a dynamic manner.

A rugged, mid-infrared (IR) CO laser absorption diagnostic has been developed to monitor the amount of CO produced by a high-pressure CH 4 -O 2 combustor test rig operating at supercritical CO 2 conditions (30 MPa and 1150°C). The laser system operates at the fundamental absorption band, ν″ = 0, R(12), of CO near 4.5 μm. The mid-IR diagnostic was constructed from a tunable quantum cascade laser (QCL), an absorption cell with two window ports for monitoring CO exhaust concentration, and two IR photodetectors. Temperature and pressure sensors were mounted near the absorption cell to monitor exhaust flow conditions, and the operational wavelength of the laser was determined by a calibration process using a known mixture of CO and N 2 . Environmental conditions at the remote outdoor test facility posed significant difficulties in the data acquisition process for the IR diagnostic. Fluctuating environmental temperatures proved to be problematic when operating cryogenic photodetectors and stabilizing a QCL designed to operate with an internal temperature of −15°C. Improvements to the IR system included elimination of problematic stagnation regions via a new absorption cell design and an increase in the CO detection limit. During steady state conditions, the mid-IR diagnostic measured the CO concentration to within ± 80.6 ppm. The IR diagnostic was shown to have superior CO detection response time and the ability to resolve features not detected by other CO detector counterparts.

Three-dimensional, three-component time-averaged velocity fields have been measured within a low-speed centrifugal fan with forward curved blades. The model investigated is representative of fans commonly used in automotive HVAC applications. The flow was analyzed at two Reynolds numbers for the same ratio of blade rotational speed to outlet flow velocity. The flow patterns inside the volute were found to have weak sensitivity to Reynolds number. A pair of counter-rotating vortices evolve circumferentially within the volute with positive and negative helicity in the upper and lower regions, respectively. Measurements have been further extended to capture phase-resolved flow features by synchronizing the data acquisition with the blade passing frequency. The mean flow field through each blade passage is presented including the jet-wake structure extending from the blade and the separation zone on the suction side of the blade leading edge.

With the onset of affordable high-performance computing and cloud computing technologies, numerical methods like computational fluid dynamics and the finite element method are spreading to all the branches of industry. The ability to perform such analyses allows engineers to speed up development and refine designs even in early stages of development without the need for expensive experimental setups. However, it is desirable to build up confidence in these results by experimental validation. GE Aviation Czech has been extensively using a variety of CFD tools to predict flow behavior in the field of internal aerodynamics — especially for design of compressor and turbine stages, combustion chambers, inlet and exhaust systems. Development of a new aerobatic engine prototype and its installation is currently under way. A series of back-to-back analyses conducted to verify performance of the aerobatic inlet and compare it with a reference commuter engine installation. The results were presented in article “GT2018-76398 Analysis and Testing of an Aerobatic Turboprop Aircraft Inlet”. A variety of critical parameters such as ram factor, pressure losses, inlet heating, and engine inlet pressure and temperature distortions were investigated. These parameters affect engine performance, operability, and compressor stability. Afterward, a unique instrumentation, data acquisition chain, and test program was developed in cooperation with the airframer. This made it possible to record all of the required parameters during ground and flight tests, including basic aerobatic maneuvers. The recorded data was then used for validation of performed CFD analyses with boundary conditions matching the conditions of ground and flight tests. Confidence in CFD results backed up by well-matched test data allows to conduct future design changes without the need to perform additional expensive experimental trials.

NASA’s Environmentally Responsible Aviation (ERA) project supported the collection of a large stereoscopic Particle Image Velocimetry (PIV) dataset of the Open Rotor Propulsion Rig (ORPR) in the 9 × 15 Low Speed Wind Tunnel at the NASA Glenn Research Center (GRC). The data collection effort acquired a volume of three component velocity measurements composed of 30 planes from near the hub radially outward towards the tip. The PIV cameras and laser were mounted to a traverse that allowed the entire data acquisition system to move from plane to plane. The PIV data acquisition was triggered on the front rotor such that the front rotor was always in the same position for each acquisition event. The aft rotor position was not recorded and varied randomly during the acquisition. Because the position of the aft rotor was not synchronized to either the forward rotor or the camera it was necessary to separate individual PIV images based off of the phase of the aft rotor before they could be processed. The phase of the aft rotor was determined by locating the outline of the rotor in the PIV images and determining its position relative to a known point. This process was conducted by an image processing algorithm. Previous algorithms were able to make a relatively accurate 3D model of the wake between the forward and aft rotors, however some small inaccuracies were present. Improvements to this algorithm allowed for more accurate phase averaging, which yielded an improved PIV dataset. Both rotors were set to the same nominal rotational speed, however variations in motor control and other physical mechanisms allowed for some differences in the true RPM of the two rotors. The effects of the aft rotor on the front rotor blade wakes within the inter-rotor flow field of the ORPR were examined. The aft rotor potential field was shown to have significant upstream impact on the front rotor wakes, altering their topology approaching the aft rotor. The wake strength was quantified through determination of the minimum velocities that occurred within a wake while a defined velocity deficit from the free stream velocity was used to determine the width of the wake. This work has potential applications to the radiated acoustics and efficiency of open rotor propulsion systems.

Several operating parameters for the control and protection of the units are acquired by the control and protection systems used in industrial applications. The use of these parameters in conjunction of physical models, empirical models and transfer functions (that represent digital replicas of the engine) allows for a broader scope of condition monitoring, taking into account the wing to wing process which spans from data acquisition to end user actionable insight. This paper describes 3 specific cases: 1) an algorithm based on the performance model of the overall GT used to monitor the axial compressor degradation and optimize the planned axial compressor water wash of an aero-derivative GT; 2) an analytic based on the flow function physic model used to monitor the clogging of the fuel nozzles in a heavy duty GT and to plan their maintenance; 3) an analytic based on a hybrid model used to monitor the axial thrust acting on a roller bearing of an aero-derivative GT and used to verify the status of the bearing and to plan its maintenance. Moreover, the paper provides details about the evaluation of the measurements, describes the model accuracy and explains how the results obtained are affected by these uncertainties and the methods used to mitigate these uncertainties. In addition, this paper shows a method to aggregate and weigh the monitoring of each single component and its own status into an overall health view.

A tip-timing system is used in a coal power station to investigate and mitigate excessive blade vibrations in the exit stage of the low-pressure steam turbine. There are presented hardware and software solutions used to monitor blade responses as well as the analyses of amplitude and frequency trends observed during the 5-year collaborative project, including operation at the nominal speed and during the shutdowns and start-ups. The transition from data acquisition to the embedded system with the partial reuse of tip-timing algorithms and LabView code is demonstrated. The proposed system processes the data coming from the turbine blades in real time and operates autonomously or under the supervision of the PC-based client program connected to the network. Acquired data are stored in a cyclic buffer and can be transferred to the host. The stack pattern is used to distinguish blades and calculate rotating reference. Tip deflection is analysed statistically and evaluated against defined reference patterns.

An engine health management (EHM) system typically consists of automated logic for data acquisition, parameter calculation, anomaly detection and eventually, fault identification (or isolation). Accurate fault isolation is pivotal to timely and cost effective maintenance but is often challenging due to limited fault symptom observability and the intricacy of reasoning with heterogeneous parameters. Traditional fault isolation methods often utilize a Single Fault Isolator (SFI) that primarily relies on gas path performance parameters. While effective for many performance related faults, such approaches often suffer from ambiguity when two or more faults have signatures that are very similar when monitored by a rather limited number of gas path sensors. In these cases, the ambiguity often has to be resolved by experienced analysts using additional information that takes many different forms, such as various non-gas path symptoms, FADEC fault codes, comparisons with the companion engine, maintenance records, and quite often, the analyst’s gas turbine domain knowledge. This paper introduces an intelligent reasoner that combines the strength of an optimal, physics-based SFI and a fuzzy expert system that mimics the analytical process of human experts for ambiguity resolution. A prototype diagnostic reasoner software has been developed and evaluated using existing flight data. Significant performance improvements were observed as compared with traditional SFI results. As a generic reasoning framework, this approach can be applied not only to traditional snapshot data, but to full flight data analytics as well.

Axial turbines for aircraft engines and power plants have reached a very high level of development. Further improvements, in particular in terms of higher efficiency and reduced number of blades and stages, resulting in higher loads, are possible, but can only be achieved through a better understanding of the flow parameters and a closer connection between experiment and numerical design and simulation. An analysis of future demands from the industry and existing turbine research rigs shows that there appears a need for a powerful turbine test rig for aerodynamic experiments. This paper deals with the development and built up of a new so called Next Generation Turbine Test Facility (NG-Turb) at the German Aerospace Center (DLR) in Göttingen. The NG-Turb is a closed-circuit, continuously running facility for aerodynamic turbine investigations, allowing independent variation of engine relevant Mach and Reynolds numbers. The flow medium (dry air) is driven by a 4-stage radial gear compressor with a high pressure ratio and a wide inlet volume flow range. In a first stage the NG-Turb test section will allow investigations on single shaft turbines up to 2½ stages. In a further expansion stage the NG-Turb will be equipped with a second independent shaft system, then enabling experiments with configurations of high and low (or intermediate) pressure turbines and in particular offering the possibility for investigations at counter rotating turbines. Secondary air for cooling investigations can be provided by auxiliary screw compressors. Mass flow through the Turbine is determined redundantly with an uncertainty of about ±0.3%, using well calibrated Venturi nozzles upstream and downstream of the test section. The operation concept and main design features of the NG-Turb will be described and an overview of the applied standard measurement and data acquisition technics capturing efficiency, traverse data etc. will be given. Thermodynamic cycle calculations have been performed in order to simulate the flow circuit of the NG-Turb and to access whether turbine operating points can be driven within the performance map of the compressor system. Finally the calibration procedure for the Venturi nozzles, which has been conducted during the commissioning phase of the NG-Turb by applying a special calibration test section, is explained and some results will be shown.

It is common to assume that the performance of low-speed turbines depend only on the flow coefficient and Reynolds number. As such the required operating point is achieved by controlling the values of these two non-dimensional quantities by, for example, appropriate choices for the mass flow rate and applied brake torque. However, when the turbine has an atmospheric inlet and uses unconditioned air, variations in ambient pressure, temperature and humidity are introduced. Whilst it is still possible to maintain the required values for the flow coefficient and Reynolds number, the ambient variations affect additional non-dimensional quantities which are related to the blade speed and gas properties. Generally, the values of these additional non-dimensional quantities cannot be controlled and, consequently, they affect the turbine performance. In addition, thermal effects, which are exacerbated by the use of plastic blades, can cause changes in the blade row seal clearance and these also affect the performance. Therefore to obtain measurements with greater accuracy and repeatability, the changes in the uncontrolled non-dimensional quantities must be accounted. This paper contains four parts. Firstly, it is described how suitable data acquisition parameters can be determined to eliminate short time scale facility unsteadiness within the measurements. Secondly, by the analysis of models, the most appropriate forms for the additional non-dimensional quantities that influence turbine performance are obtained. Since the variations in the uncontrolled non-dimensional quantities affect repeatability the size of the effect on the turbine performance is quantified. Thirdly, a best-fit accounting methodology is described which reduces the effects of the uncontrolled non-dimensional quantities on turbine performance provided sufficient directly related measurements are available. Finally, the observations are generalised to high-speed turbomachines.

In contrast to other alternative fuels, methanol offers a plethora of advantages; it can be produced from natural gas, coal or any organic biomass matter. Methanol is the cheapest of all alternative fuels and has a reduced likelihood of reacting to form pollutants such as soot, CO, NOx. Furthermore, methanol is a liquid at standard atmospheric conditions and can be stored and transported much more cheaply and safely than gaseous fuel. There are also drawbacks that impinge on its overall utility: Methanol has a lower energy density as compared to conventional fossil liquid fuels (kerosene, diesel oil etc.), it has the propensity to react to form aldehyde emissions and to dissolve in atmospheric moisture. Methanol also has a low vapor pressure that could lead to cold start problems in IC engines, at temperatures below 15°C. The objective of the present work is to perform an experimental investigation and a chemical kinetic comparison of combustion characteristics between methanol and kerosene & diesel fueled swirl stabilized burner. A special emphasis is given to the impact of fuel conversion of existing combustion systems from kerosene (or diesel) to methanol. The experimental set-up is based on a modified industrial burner with heating power of 50kW. In order to have a better control over the incoming air flow, the original blower is replaced with a larger and more stable industrial blower that allowed precise monitoring of the air flow rate. The combustion chamber consists of a stainless steel tube, 10cm in diameter and 80cm in length. The experiments include measurements of temperature distribution inside the combustor: wall temperature distribution is recorded with the use of thermocouples and a calibrated infra-red camera. The composition of the combustion pollutants is also monitored at the exhaust. A National Instruments cRIO 9074 controller is used in conjunction with a National Instruments LabView interface for data acquisition. The comparison between methanol and kerosene characteristics is carried out for equal heat release and equivalence ratios. Experimental results show that methanol burns slower than kerosene and therefore requires a longer combustor. It is found that the larger methanol droplet size and its larger volume flux contribute significantly to this extended length requirement for complete methanol combustion. The measured CO emission values for kerosene and methanol were 25 and 110 ppm respectively, and 40 and 10ppm for NOx. These results clearly indicate the reduced NOx emission during methanol combustion; however, the notable presence of CO indicates that methanol needs a longer combustor length to complete the combustion process.

The present work is part of a comprehensive heat transfer and film-cooling study on a locally cooled non-axisymmetric contoured turbine endwall. A new test rig consisting of a linear cascade of three prismatic vanes at unity scale and exchangeable endwall has been established. The rig is operated in an open-loop configuration at a reduced main gas temperature of 425 K, an exit Mach number of 0.5 and an exit Reynolds number of 1.6×10 6 . Air is used both as main gas and coolant; a realistic density ratio is achieved by cooling the coolant below freezing. In Part I [1] of the study aerodynamic measurements are presented. This paper concentrates on film-cooling of the contoured endwall with special emphasis on data acquisition and reduction for the application of the superposition principle of film cooling. The first experimental results from thermographic measurements are discussed.

An ambitious project in propulsion was introduced as part of the final-year integrator project offerings of the mechanical and aerospace engineering programs at École Polytechnique de Montréal in 2011–2012. It has been running successfully for the past three academic years. The project consists in the design, fabrication and placement into service of a functional instrumented multi-stage compressor test rig, including the compressor, for research in compressor aerodynamics. A team of 15–17 senior-year undergraduate engineering students are given set of design and performance specifications and measurement requirements, an electric motor and drive, a data acquisition system and some measurement probes. They must complete the project in two semesters with a budget on the order of Can$15,000. The compressor is made from rapid prototyping to keep production cost and time reasonable. However, its required rotation speed of 7200 rpm stretches the limits of the plastic material and presents the same structural challenges as industrial compressors running at higher speeds. The students are split into sub-teams according to the required disciplines, namely compressor aerodynamics, general aerodynamics, structures, dynamics, mechanical design and integration, instrumentation and project management. For the initial phase, which covers the first two months, the students receive short seminars from experts in academia and industry in each discipline and use the knowledge from fundamental engineering courses to analytically model the different components to come up with a preliminary design. In the second phase, covering months three through six, the students are trained at commercial simulation tools and use them for detailed analysis to refine and finalize the design. In each of the first two phases, the students present their work in design reviews with a jury made up of engineers from industry and supervising professors. During the final phase, the compressor is built and tested with data acquisition and motor control programs written by the students. Finally, the students present their results with comparison of measured performance with numerical and analytical predictions from the first two phases and hand over their compressor rig with design and test reports as well as a user manual and an assembly/maintenance manual. This complete project allows the students to put into practice virtually all the courses of their undergraduate engineering curriculum while giving them an extensive taste of the rich and intellectually challenging environment of gas turbine and turbomachinery engineering.

A gas turbine generator set intended for use as an emergency power source has been evaluated and tested using a computer controlled data acquisition and processing system. At 1-sec intervals during startup, transducer devices monitor such items as turbine inlet temperature, compressor discharge pressure and speed. While operating under normal or steady-state conditions, a total of 65 parameters are logged at 5-min intervals. In addition, power, specific fuel consumption and efficiency are calculated every 5 min. Computer output is in engineering units (deg F, psi, cfm, etc.). The computer programming language is structured with Fortran-like arithmetic and formatting capabilities. Analog-to-digital conversion is accomplished by an integrating digital voltmeter and switching via a 200-channel reed relay scanner. Two teletypewriters, a programmable clock, and 64,000 words of disk storage comprise the rest of the major hardware. For this project 4096 words of computer core were implemented. A turbine development program conducted with this computerized data system can be completed by a single appropriately trained individual in approximately half the time required by a team of three or four using conventional data gathering methods.

The Engine Test Facility of the National Gas Turbine Establishment has used computer-controlled instrumentation systems for on-line data acquisition and processing for over 15 yr. The development of the system is outlined and the current system described. To illustrate its present use typical results obtained during tests on a large turbofan engine are presented and discussed. The emphasis of this paper is on the use of the system to assess the accuracy and reliability of the test measurements before commencing qualification tests on the engine. The probable sources of error are described and placed into three categories, to assist handling and propagation through the calculation processes. The resultant uncertainty in measurement of Thrust and sfc at cruise (M = 0.85, 35,000 ft), is estimated to be 0.50 percent (one standard deviation) for a single performance curve.

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.

Improvements of large steam turbines require results from experimental investigations on real turbines. At the Institute fur Thermische Stromungsmaschinen a new Last-Stage test stand has been built providing extended facilities for investigations of flow, steam moisture and vibration phenomona. This contribution presents a description of the test stand with its various set-up configurations for the evaluation of the data to be obtained. The measurement, data acquisition and evaluation systems are explained.