Boundary Layer Ingestion (BLI) is a propulsion technology being investigated at NASA by the Advanced Aircraft Transportation Technology (AATT) Program to facilitate a substantial reduction in aircraft fuel burn. In an attempt to experimentally demonstrate an increase in propulsive efficiency of a BLI engine, a first-of-its-kind sub-scale high-bypass ratio 22” titanium fan, designed to structurally withstand significant unsteady pressure loading caused by a heavily distorted axial air inflow, was built and then tested in the transonic section of the GRC 8′ × 6′ Supersonic Wind Tunnel. The vibratory responses of a subset of fan blades were measured using strain gages placed in four different blade pressure side surface locations. Response highlights include a significant response of the blade’s first resonance to engine order excitation below idle as the fan was spooled up and down. The fan fluttered at the design speed under off operating line, low flow conditions. This paper presents the blade vibration response characteristics over the operating range of the fan and compares them to predicted behaviors. It also provides an assessment of this distortion tolerant fan’s (DTF) ability to withstand the harsh dynamic BLI environment over an entire design life of billions of load cycles at design speed.

Incorporating recuperators into turbine engines which enables the heat transfer between exhaust gas and compressed air, indicates considerable potential for lower emissions and SFC (specific fuel consumption). However as a matter of fact, they have not yet found wide acceptance in aircraft applications. One of the main potential disadvantages is because in such situations system overall weight needs to be strictly controlled, and there are concerns that the beneficial fuel saving may not offset the additional bulk and weight of the recuperator. In this work, aimed at evaluating the performance of the recuperated rotorcraft and investigating the influence of a recupeator on the whole system, a comprehensive simulation framework has been developed which mainly contains three modules for various flight conditions: Helicopter performance module (HPM), GasTurb engine simulation module and the recuperator weight estimation. Specifically, HPM calculates the helicopter power requirement for different flight conditions (altitude, speed, etc.), and GasTurb is used to simulate the performance of conventional and recuperated turbine engine, and compute the corresponding engine operating point to meet the power demand. The recuperator weight estimation is mainly based on previous studies as a function of mass flow and effectiveness. The methodology is adopted for a typical twin engine light helicopter configuration. In comparison with the conventional non-recuperated cycle, the authors studied the fuel saving potential of the recuperated cycle for different recuperator effectiveness under various flight conditions, and a trade-off analysis was also conducted to identify the flight time required to compensate the additional recuperator weight. The obtained results suggested that the recuperated cycle possesses great potential, especially for long duration and large range mission, but it may not be necessarily suitable for all types of helicopter missions.

Aircraft engine maintenance is performed on an on-condition basis. Monitoring the engine condition during operation is important to provide an efficient maintenance. Engine Condition Monitoring has thus become a standard procedure during operation. However, one of the most important parameters, the engine thrust, is not directly measured and can therefore not be monitored, which makes it difficult to distinguish whether deteriorating trends e.g. in fuel comsumption must be attributed to the engine (e.g. due to thermodynamic wear) or to the aircraft (e.g. due to increased drag). Being able to make this distinction would improve troubleshooting and maintenance planning and thus help to reduce the cost of ownership of an aircraft. This paper describes the development and quality assessment of a system for direct engine thrust measurement during the normal engine operation. The system was designed, calibrated and validated with engine test runs. After the necessary certification of the whole system a flight test campaign to validate the system, when installed on an aircraft, was started. In the presented work an assessment of the quality of measured data from the first period of the ongoing flight test is presented.

Central targets for jet engine research activities comprise the evaluation of improved engine components and the assessment of novel engine concepts for enhanced overall engine performance in order to reduce the fuel consumption and emissions of future aircraft. Since CO 2 emissions are directly related to engine fuel burn, a reduction in fuel consumption leads to lower CO 2 emissions. Therefore improvements in engine technologies are still significant and a multi-disciplinary pre-design approach is essential in order to address all requirements and constraints associated with different engine concepts. Furthermore, an increase in effectiveness of the preliminary design process helps reduce the immense costs of the overall engine development. Within the DLR project PEGASUS (Preliminary Gas Turbine Assessment and Sizing) a multi-disciplinary pre-design and assessment competence of the DLR regarding aero engines and gas turbines was established. The application of modern preliminary design methods allows for the construction and evaluation of innovative next generation engine concepts. The purpose of this paper is to present the developed multi-disciplinary pre-design process and its application to three aero engine models. First, a state of the art twin spool mixed flow turbofan engine model is created for validation purposes. The second and third engine models investigated comprise future engine concepts: a Counter Rotating Open Rotor and an Ultra High Bypass Turbofan. The turbofan used for validation is based on publicly available reference data from manufacturing and emission certification. At first the identified interfaces and constraints of the entire pre-design process are presented. An important factor of complexity in this highly iterative procedure is the intricate data flow, as well as the extensive amount of data transferred between all involved disciplines and among the different fidelity levels applied within the smoothly connected design phases. To cope with the inherent complexity data modeling techniques have been applied to explicitly determine the required data structures of those complex systems. The resulting data model characterizing the components of a gas turbine and their relationships in the design process is presented in detail. Based on the established data model the entire engine pre-design process is presented. Starting with the definition of a flight mission scenario and the resulting top level engine requirements thermodynamic engine performance models are developed. By means of these thermodynamic models, a detailed engine component pre-design is conducted. The aerodynamic and structural design of the engine components are executed using a stepwise increase in level of detail and are continuously evaluated in the context of the overall engine system.

In the present paper, the synergistic combination of inter-cooling with pulsed detonation combustion is analyzed concerning its contribution to NO x and CO 2 emissions. CO 2 is directly proportional to fuel burn and can, therefore, be reduced by improving specific fuel consumption and reducing engine weight and nacelle drag. A model predicting NO x generation per unit of fuel for pulsed detonation combustors, operating with jet-A fuel, is developed and integrated within Chalmers University’s gas turbine simulation tool GESTPAN. The model is constructed using CFD data obtained for different combustor inlet pressure, temperature and equivalence ratio levels. The NO x model supports the quantification of the trade-off between CO 2 and NO x emissions in a 2050 geared turbofan architecture incorporating intercooling and pulsed detonation combustion and operating at pressures and temperatures of interest in gas turbine technology for aero-engine civil applications.

Efforts to achieve NASA’s N+2 and N+3 fuel burn goals have led to various future aircraft concepts. A commonality in all these concepts is the presence of a high degree of interaction among the various disciplines involved. A tightly integrated propulsion/airframe results in distortion in the flow field around the engine annulus. Although beneficial in terms of propulsive efficiency (due to boundary layer ingestion), the impact of distortion on fan performance and operability remains in question for these concepts. As such, rapid evaluation of the impacts of distortion during the conceptual design phase is necessary to assess various concepts. This is especially important given the expansion of the design space afforded by turbo-electric and hybrid-electric distributed propulsion concepts, in which the gas turbine generator and propulsive devices can be decoupled in space. A simple and rapid methodology to assess operability of compressors is the theory of Parallel Compressors (PC). PC theory views the compressor as two compressors in parallel, one with a uniform high P t and the other with a uniform low P t , both operating at the same speed and exiting to a common static pressure. The assumption of two compressors exiting at the common static pressure is not entirely true, especially when the distortion is high. In this paper, the development of a modified parallel compressor model with parametric boundary condition that can capture the impact of non-uniform inflow on fan performance is introduced and validated. Unlike classical PC model, the modified approach introduces a boundary condition dependent on the intensity of distortion (DPCP) at the Aerodynamic Interface Plane (AIP). Additionally, the concept of PC is also extended to Multi-Per Revolution (MPR) distortion. A modeling environment which follows this methodology is created in PROOSIS, an object oriented 0-D cycle code. The model was created using the “compressor” components acting in parallel and a procedure for implementing both design mode and off-design mode solutions was created using the PROOSIS toolset. The example problem was implemented to demonstrate two capabilities — i) the ability of quantifying impacts on thrust and performance of a ducted fan propulsion system, and ii) the ability of predicting loss in stability pressure ratio. The results clearly show the ability of the tool to quantify distortion related losses. The work described in this paper can be integrated to a Multi-Disciplinary Design and Optimization (MDAO) framework along with other disciplines and can be used to evaluate the viability of design space offered by novel aircraft configurations.

Gas turbines in general and aircraft engines in particular undergo frequently dynamic operations. These operations include the routine start-up, load change and shut downs to cover their operation envelope. The frequency of the dynamic operation depends on the size of the engines and the field of application. Engines for commuter aircrafts and particularly helicopter engines operate more often in an off-design mode compared to large commercial aircraft engines and power generation gas turbines. During these routine operations, the compressor mass flow, the pressure ratio, the combustion chamber fuel and air mass flow as well as turbine mass flow change. These changes affect the engine aerodynamic performance and its efficiency. To avoid the inception of rotating stall and surge, high performance gas turbines are equipped with mechanisms that adjust the stator stagger angles thus aligning the stator exit flow angle to the rotor inlet angle, which reduces an excessive incidence. The reduction of incidence angle not only preserves the stable operation of the compressor but it also prevents the compressor efficiency from deterioration. The existence of an inherent positive pressure gradient may cause the boundary layer separation on compressor blades leading to the rotating stall and surge. Such condition, however, does not exist in a turbine, and therefore, there has been no compelling reason to apply the blade adjusting method to the turbine component. For the first time, the impact of turbine blade stagger angle adjustment on the gas turbine efficiency during the operation is shown in this paper. Given a statistically distributed load condition, the extensive dynamic simulation reported in this paper shows how the efficiency can be positively affected through proper blade adjustment. For the time dependent operation, the code GETRAN developed by the author was enhanced to include the turbine blade adjustment as a function of time. To conduct the dynamic simulation with turbine stator stagger angle adjustment during a dynamic operation, the full geometry of the Brown Boveri GT-9 gas turbine was utilized. Starting from the reference stagger angle, it is varied within an incidence range of ± 3 degree. Detailed simulation results show the substantial efficiency improvement through stator stagger blade adjustment.

The MEE (More Electric Engine) is a concept for engine system electrification and is an evolutionary step in engine system design that contributes to the reduction of aviation CO 2 emissions. Mifee (Metering and integrated fuel feeding electrification) and the E3M (Engine Embedded Electric Machine) are the key technologies of the MEE. The purpose of engine thermal management is maintaining the balance between heat generation by the engine system and heat dissipation to the outside of the engine. In recent engine system designs, thermal system design has become an issue because of increased heat generation within the system. For example, a recently developed turbo-fan engine system increases the heat generation by introduction of a fan drive gear system that produces a large amount of heat in addition to the conventional heat source, such as engine main bearings and gears. The MEE will have further heat sources within its system, like the E3M, which is a high-power electric machine. In this paper, an investigation approach and the result of a feasibility study of the MEE thermal management system is described. In addition, the perspective of the technology trend from the MEE toward future hybrid propulsion is also discussed. The global requirements for climate protection strongly demand game-changing technology that significantly improves the aircraft’s overall efficiency. A series/parallel partial hybrid propulsion system, in which both a turbo-fan engine and electrical motor-driven fans generate propulsive power, is considered to be one of the most promising approaches for the future commercial aircraft hybrid propulsion system. The MEE and E3M technology evolves until it will be applied in hybrid propulsion system.

In order to decrease the fuel consumption, a new flight mode is being considered for twin-engine helicopters, in which one engine is put into sleeping mode (a mode in which the gas generator is kept at a stabilized, sub-idle speed by means of an electric motor, with no combustion), while the remaining engine operates at nominal load. The restart of the engine in sleeping mode is therefore deemed critical for safety reasons. This efficient new flight mode has raised the interest in the modeling of the restart of a turboshaft engine. In this context, the initial conditions of the simulations are better known relative to a ground start, in particular the air flow through the gas generator is constant, the fuel and oil system states are known and temperatures of the casings are equal to ambient. During the restart phase of the engine, the gas generator speed is kept at constant speed until the light-up is detected by a rise in inter-turbine temperature, then the starter torque increases, accelerating the engine towards idle speed. In this paper, the modeling of the acceleration of the gas generator from light-up to idle and above idle speeds is presented. Details on the light-up process are not addressed here. The study is based on the high-fidelity aero-thermodynamic restart model that is currently being developed for a 2000 horse power, free turbine turboshaft. In this case, the term high-fidelity refers not only to the modeling of the flow path components but it also includes all the subsystems, secondary air flows and controls with a high level of detail. The physical phenomena governing the acceleration of the turboshaft engine following a restart — mainly the transient evolution of the combustion efficiency and the power loss by heat soakage — are discussed in this paper and modeling solutions are presented. The results of the simulations are compared to engine test data, highlighting that the studied phenomena have an impact on the acceleration of the turboshaft engine and that the model is able to correctly predict acceleration trends.

Detailed information of the thermodynamic parameters, system performance and operating behavior of aircraft APU cycles is rarely available in literature. In order to set up numeric models and study cycle modifications, validation data with well defined boundary conditions is needed. Thus, the paper introduces an APU test rig based on a Garrett GTCP36-28 with detailed instrumentation which will be used in a further step as a demonstration platform for cycle modifications. The system is characterized in the complete feasible operating range by alternating bleed air load and electric power output. Furthermore, simulations of a validated numerical cycle model are utilized to predict the load points in the operating region which were unstable during measurements. The paper reports and discusses turbine shaft speed, compressor air mass flow, fuel mass flow, efficiencies, compressor outlet pressure and temperature, turbine inlet and outlet temperature as well as exhaust gas emissions. Furthermore, the results are discussed with respect of the difference compared to a Hamilton Sundstrand APS3200. Though the efficiencies of the GTCP36-28 are lower compared to the APS3200, the general behavior is in good agreement. In particular, the effects of separate compressors for load and power section are discussed in contrast to the GTCP36-28 system design comprising a single compressor. In general, it was shown that the GTCP36-28 is still appropriate for the utilization as a demonstration platform for cycle modification studies.

This paper will discuss the development and testing of an electric drive option designed for the propulsion system of the US Navy’s DDG 51 Class ships. It will briefly explain the history of the Hybrid Electric Drive (HED) program, including that of its predecessor, Proof of Concept (PoC), and the HED’s planned shipboard installation schedule. Operating at lower ship speeds, in a range where the currently installed propulsion gas turbines are less fuel efficient, the HED is expected to increase the ship’s fuel economy, allowing the ship to remain on station accomplishing its mission for a longer period of time. This paper will discuss how the gas turbine propulsion system, in concert with the HED, will be used to provide the most fuel efficient drive combination for various operating scenarios. Also covered will be a description of the major stakeholders involved in the HED’s development and implementation along with some of the constraints and challenges that were encountered in the testing phase of the program, both at the OEM facilities and at the US Navy’s Land Based Engineering Site (LBES) in Philadelphia PA. Planned fuel economy testing results obtained at the LBES facility will also be presented, intended to determine an estimate of the fuel savings that can be expected when the system is first placed in service on USS TRUXTUN (DDG 103) July 2018.

The article describes a refining method for a fuel pump of rocket powerful turbo-pump unit by the joint usage of mathematical optimization software IOSO, meshing complex NUMECA and CFD complex ANSYS CFX. The optimization software was used for automatic change of the geometry of low-pressure impeller, transition duct and high-pressure impeller to find the optimal design. It was mandatory to keep the original variant of the remaining parts of the pump. For this reason, only geometrical parameters of the blades were varied without changing the contours of the pump meridional flow part. The investigated pump consists of five parts: inlet duct, low-pressure screw centrifugal stage, transition duct, high-pressure screw centrifugal stage and volute outlet duct. The pump main parameters with water as the working fluid (based on experiment data) were the following: high-pressure stage rotor speed was 13300 rpm; low-pressure rotor speed was 3617 rpm by gearbox; inlet total pressure was 0.4 MPa; outlet mass flow was 132.6 kg/s at the nominal mode. Creation of vane unit mesh (rotors and stator transition duct) was performed using NUMECA AutoGrid5. Sector models were used for the calculation simplification. The flow around only one blade or screw was considered. Setting up and solution of the task were carried out in the ANSYS CFX solver. Comparison of calculated characteristics of the basic pump with the experimental data was performed before the optimization. The analysis of characteristics for the obtained optimized pump geometry was carried out. It was found that pump with optimized geometry has greater efficiency in comparison with the original pump variant. The obtained reserve can be used to boost the rocket engine, and/or to reduce the loading of the main turbine, which operates in aggressive oxidizing environment.

The development of new generations of aircraft engines with reduced environmental impact heavily relies on high-fidelity 3D numerical analysis of the main engine components, compressor, combustion chamber, turbine and their interactions, including the transient and off-design behavior of the full engine. Unlike component-by-component analysis, which requires separate assumptions for the pressure and temperature boundary conditions for each component, a fully coupled approach requires only knowledge of the compressor inlet and turbine outlet flow conditions. In addition, the engine rotation speed can also be varied during the simulation to converge to the correct balance of power between compressor and turbine. This integrated approach provides a detailed description of the flow field inside the full engine at the desired operating point with one single CFD simulation. The full engine simulation methodology can be developed at several levels: (1) RANS simulations with mixing-plane interfaces between components; (2) advanced RANS treatment with inputs from the nonlinear harmonic (NLH) methodology to allow for tangential non-uniformity, such as hot streaks entering the turbine nozzle from the combustor; (3) inclusion of the unsteady rotor-stator interactions, via NLH, in compressor and turbine stages; (4) coupling with LES simulations in the combustor. This paper presents results from levels (1) and (2) of this methodology applied to a micro-turbine gas engine including the HP compressor, combustor, HP and LP turbines and the exhaust hood. The geometry has been obtained from the redesign of the KJ66 micro gas turbine engine using preliminary design tools. The injection and burning of fuel inside the combustion chamber are modeled with a simplified flamelet model. The paper presents the approach and results of the full engine simulation; as well as the initial steps towards level (3).

Fouling affects gas turbine operation and airborne or fuel contaminants, under certain conditions, become very likely to adhere to surfaces if impact takes place. Particle sticking implies the change in shape in terms of roughness of the impinged surface. The consequences of these deposits could be dramatic: these effects can shut an aircraft engine down or derate a land-based power unit. This occurrence may happen due to the reduction of the compressor flow rate and the turbine capacity, caused by a variation in the HPT nozzle throat area (geometric blockage due to the thickness of the deposited layer and the aerodynamic blockage due to the increased roughness, and in turn boundary layer). Several methods to quantify particle sticking have been proposed in literature so far, and the experimental data used for their validation vary in a wide range of materials and conditions. The experimental analyses have been supported by (and have given inspiration to) increasingly realistic mathematical models. Experimental tests have been carried out on (i) a full scale gas turbine unit, (ii) wind tunnel testing or hot gas facilities using stationary cascades, able to reproduce the same conditions of gas turbine nozzle operation and finally, (iii) wind tunnel testing or hot gas facilities using a coupon as the target. In this review, the whole variety of experimental tests performed is gathered and classified according to composition, size, temperature and particle impact velocity. Using particle viscosity and sticking prediction models, over seventy (70) tests are compared with each other and with the model previsions providing a useful starting point for a comprehensive critical analysis. Due to the variety of test conditions, the related results are difficult to be pieced together due to differences in particle material and properties. The historical data of particle deposition obtained over thirty (30) years are classified using particle kinetic energy and the ratio between particle temperature and its softening temperature. Qualitative thresholds for the distinction between particle deposition, surface erosion and particle break-up, based on particle properties and impact conditions, are identified. The outcome of this paper can be used for further development of sticking models or as a starting point for new insight into the problem.

High velocity oxygen fuel (HVOF) sprayed stellite coating is often adopted on contact surfaces of the shrouds and stubs of titanium turbine blades, in order to increase fretting fatigue strength and reduce fretting wear. To confirm the effectiveness of the sprayed stellite coating, we conducted friction-type fretting tests originally developed to simulate the load condition of shrouds and stubs, vibratory force was carried only by friction force. We also investigated the fretting fatigue mechanism of coated materials by observing non-propagating cracks in the tested specimens and analyzed crack propagation behavior at the interface using fracture mechanics. The fretting fatigue tests confirmed that the sprayed stellite coating could double the fretting fatigue strength of Ti-6A1-4V. The fretting fatigue mechanism of Ti-6A1-4V with a sprayed stellite coating was revealed as follows: Final fracture was led by a newly initiated micro-crack at the interface as cracks initiated on the coating surface did not continuously propagate into the substrate but mainly propagated along the interface. The micro-crack propagation at the substrate depends on the length of interface crack. The longer interface crack delays micro-crack propagation to the substrate because the lines of force are less concentrated at the crack tip. Hence, the interface crack makes the fretting fatigue strength of coated Ti-6A1-4V higher than that of uncoated specimens. This finding was supported by the cross-sectional observations of non-propagating cracks in the tested specimens, and by FE analysis which revealed that Δ K of the substrate crack initiated at the interface decreases with an increase in delaminated crack length. Ti-6A1-4V with a sprayed stellite coating offers better wear resistance than uncoated Ti-6A1-4V because the coating’s friction coefficient is lower than that of Ti-6Al-4V, although the wear rates of coated and uncoated materials against the consumption energy per cycle were almost the same.

In this study, research activity concerning the utilization of biogas into a commercial micro gas turbine (MGT) Ansaldo Turbec T100P is presented. The use of biogas in MGT has been already quite explored in literature. However, conflicting and not yet fully comprehensive results highlight the need for further scientific investigations. Although several authors have demonstrated, through numerical studies, the possibility of using biogas with amounts of CO 2 up to 40–50% in volume in commercial MGT, experimental evidences of this condition are not yet present in literature. In fact, the CO 2 concentration limit experimentally validated in literature, after which phenomena of detachment and instability of the flame happen, is considerably lower (∼12%). The utilization of biogas in micro gas turbines generally requires structural adaptations, often invasive and expensive, to the fuel compression unit and control valves. Not infrequently a partial redesign of the combustor is often necessary. Costs associated with the modifications significantly reduce the benefits of a rational and sustainable exploitation of energy resources of lower value as the biogas. Aim of this activity is to overcome these issues, by identifying a non-invasive and cost-effective technical solution to increase the CO 2 concentration in the fuel and to extend the operation domain when low calorific value fuels are used. It has been studied, implemented and tested an innovative management strategy of fuel feeding and combustion. Experimental layout and preliminary results of this novel approach are shown and discussed.

The purpose of the paper is the investigation of the combustion development in a Lean Premixed Prevaporized (LPP) combustor supplied with liquid fuels of a Micro Gas Turbine. The injection system is equipped by an airblast atomizer with three air inlets. In such a study the analysis is performed by a CFD tool that can simulate the injection conditions, by isolating and studying some specific phenomena. A 3-D fluid dynamic code (the FLUENT ® flow solver) has been used to simulate the spray pattern in the chamber fuelled by kerosene fuel. Preliminarily, the numerical simulation refers to cold flow conditions, in order to validate the estimation of the fundamental spray parameters, such as the spray angle and Sauter Mean Diameter of the droplets; in a second step, the calculations employ boundary conditions close to those occurring in the actual combustor operation, in order to predict the fuel vapour distribution throughout the premixing chamber. In particular, the fuel is injected under the typical conditions that occur in the injection system of a gas turbine LPP combustor and in all cases examined, the boundary information is introduced in terms of air and fuel mass flow rates and of inlet characteristics of the air flow entering the revalorizing chamber, in order to estimate the fuel vapour formation and its distribution. In a third phase the combustion phenomenon is simulated and the NO emissions calculated. Finally, the best solution obtained is also tested by using a bio-fuel to compare the combustion performance and NO amount.

Ammonia is expected to be a hydrogen carrier that has potential as a carbon-free fuel. Ammonia is known as a nonignitable fuel, and it is not easy to hold ammonia flames under atmospheric conditions. A demonstration test with the aim of showing the potential of ammonia-fired power plants was conducted using a micro gas turbine. A 50-kW-class turbine system firing kerosene was selected as a base model. More than 40 kW of power generation was achieved by firing ammonia gas or a mixture of ammonia and methane by modifying the combustor, the fuel control device, and the gas turbine startup sequence. The prototype bifuel combustor is a swirl combustor employing a non-premixed flame and a decreased air flow rate near a gas fuel injector for flame holding. Ammonia combustion in the prototype bifuel combustor was enhanced by supplying hot combustion air and by modifying the air inlets. However, the exhaust gases from the ammonia flames had high NOx concentrations. NOx removal equipment using selective catalytic reduction can reduce NOx emission levels to below 10 ppm from more than 1000 ppm (converted value of NOx to 15% O 2 ) as already reported. However, downsizing of NOx removal equipment should be achieved for practical use. Therefore, a low NOx combustor was developed. As the first step of the development of the combustor, flame observation in the gas turbine combustor was tried. Although the observation area was limited, an inhomogeneous swirling orange flame of ammonia gas was observed. Then, a combustor test rig was prepared for a detailed observation of ammonia flame under various combustion conditions. The combustor test rig used a regenerative heat exchanger for heating combustion air, and it used an orifice for pressure drop instead of a turbine. Combustion air and cooling air were supplied from two air compressors. At startup of the combustor test rig, a spark plug was used to ignite non-premixed methane and air. After heating the regenerative heat exchanger, ammonia gas was supplied to the combustor instead of methane gas. The exhaust gases from the combustor were analyzed using FTIR (Fourier transform infrared spectroscopy) under various conditions, such as methane firing, methane–ammonia firing, and ammonia firing. Although there are several concepts for NOx reduction, a rich–lean combustion method was applied first for ammonia firing. The rich–lean combustor modified from the prototype bifuel combustor also could burn ammonia well in cases of both methane–ammonia firing and ammonia firing. The rich–lean combustor succeeded in reducing NOx emission from methane–ammonia combustion to half the value measured in the case of the prototype bifuel combustor.

Modern lean burn aero-engine combustors make use of relevant swirl degrees for flame stabilization. Moreover important temperature distortions are generated, in tangential and radial directions, due to discrete fuel injection and liner cooling flows respectively. At the same time, more efficient devices are employed for liner cooling and a less intense mixing with the mainstream occurs. As a result, aggressive swirl fields, high turbulence intensities and strong hot streaks are achieved at the turbine inlet. In order to understand combustor-turbine flow field interactions, it is mandatory to collect reliable experimental data at representative flow conditions. While the separated effects of temperature, swirl and turbulence on the first turbine stage have been widely investigated, reduced experimental data is available when it comes to consider all these factors together. In this perspective, an annular three-sector combustor simulator with fully cooled high pressure vanes has been designed and installed at the THT Lab of University of Florence. The test rig is equipped with three axial swirlers, effusion cooled liners and six film cooled high pressure vanes passages, for a vortex-to-vane count ratio of 1:2. The relative clocking position between swirlers and vanes has been chosen in order to have the leading edge of the central NGV aligned with the central swirler. In order to generate representative conditions, a heated mainstream passes though the axial swirlers of the combustor simulator, while the effusion cooled liners are fed by air at ambient temperature. The resulting flow field exiting from the combustor simulator and approaching the cooled vane can be considered representative of a modern Lean Burn aero engine combustor with swirl angles above ±50°, turbulence intensities up to about 28% and maximum-to-minimum temperature ratio of about 1.25. With the final aim of investigating the hot streaks evolution through the cooled high pressure vane, the mean aerothermal field (temperature, pressure and velocity fields) has been evaluated by means of a five hole probe equipped with a thermocouple and traversed upstream and downstream of the NGV cascade.

In modern lean burn aero-engine combustors, highly swirling flow structures are adopted to control the fuel-air mixing and to provide the correct flame stabilization mechanisms. Aggressive swirl fields and high turbulence intensities are hence expected in the combustor-turbine interface. Moreover, to maximize the engine cycle efficiency, an accurate design of the high pressure nozzle cooling system must be pursued: in a film cooled nozzle the air taken from last compressor stages is ejected through discrete holes drilled on vane surfaces to provide a cold layer between hot gases and turbine components. In this context, the interactions between the swirling combustor outflow and the vane film cooling flows play a major role in the definition of a well performing cooling scheme, demanding for experimental campaigns at representative flow conditions. An annular three-sector combustor simulator with fully cooled high pressure vanes has been designed and installed at THT Lab of University of Florence. The test rig is equipped with three axial swirlers, effusion cooled liners and six film cooled high pressure vanes passages, for a vortex-to-vane count ratio of 1:2. The relative clocking position between swirlers and vanes has been chosen in order to have the leading edge of the central airfoil aligned with the central swirler. In this experimental work, adiabatic film effectiveness measurements have been carried out in the central sector vanes, in order to characterize the film-cooling performance under swirling inflow conditions. The Pressure Sensitive Paint technique, based on heat and mass transfer analogy, has been exploited to catch highly detailed 2D distributions. Carbon dioxide has been used as coolant in order to reach a coolant-to-mainstream density ratio of 1.5. Turbulence and five hole probe measurements at inlet/outlet of the cascade have been carried out as well, in order to highlight the characteristics of the flow field passing through the cascade and to provide precise boundary conditions. Results have shown a relevant effect of the swirling mainflow on the film cooling behaviour. Differences have been found between the central airfoil and the adjacent ones, both in terms of leading edge stagnation point position and of pressure and suction side film coverage characteristics.