bluuebunny/crossref_metadata_2025_split huggingface dataset:
Selection and Development of the World’s Most Power-Dense Gas Turbine Module for the New Korean Frigate
Authors: Lim Sangpil , Harvey Adam 
Published: September 2016

The MT30 marine gas turbine has been developed specifically for 21st century naval propulsion using modern techniques and methods. Design and development of the MT30 began in 1999 and has since been qualified for naval service following extensive testing. Since then the engine has rapidly been adopted by progressive navies, in both its mechanical and electrical power generation configuration.
The Lockheed Martin Littoral Combat Ship (LCS) is one of a new class of United States Navy (USN) fast combatants which has been at sea for more than six years and is powered by the MT30. A combined MT30-driven generator was selected for the new USN DDG1000 Zumwalt class of destroyer and has also been successfully installed into the Royal Navy’s Queen Elizabeth Class aircraft carrier. Most recently, the MT30 Compact Package has been selected to power the Royal Navy’s Type 26 Global Combat Ship which will be built by BAE Systems.
The MT30 Compact Package has been designed with the aim of powering modern warship programmes, with the result that it is currently the World’s most power dense in-service marine gas turbine. This is an important factor in naval propulsion where delivering a high power output in a compact space is essential.
In addition to the programmes stated above, the MT30 Compact Package was selected for the new Republic of Korea Navy’s (RoKN) frigate programme with a single-GT CODLOG hybrid arrangement consisting of propulsion motors and a Diesel-electric system.
As a result, Rolls Royce was selected by the RoKN to deliver the MT30 Gas Turbine Unit and, from a preliminary Rolls-Royce compact package design, the engine and machinery division of Hyundai Heavy Industry (HHI-EMD) has developed the Compact Package for the New Korea Frigate. The MT30 GT was delivered to the HHI-EMD facility in 2014 with the surrounding Compact Package built at HHI-EMD before onward delivery to Daewoo Shipbuilding and Marine Engineering (DSME) where construction of the first frigate will take place.
This paper provides the rationale for selection of the MT30 Compact Package for the New Korea Frigate Programme and also describes the development of the MT30 Compact Package; aspects of the design process, construction of the Compact Package and the factory acceptance test conducted at the HHI-EMD facility.
Design of an Intermediate Turbine Duct Downstream of a High Pressure Turbine
Authors: Hou Chaoshan , Wu Hu 
Published: September 2016

The flow leaving the high pressure turbine should be guided to the low pressure turbine by an annular diffuser, which is called as the intermediate turbine duct. Flow separation, which would result in secondary flow and cause great flow loss, is easily induced by the negative pressure gradient inside the duct. And such non-uniform flow field would also affect the inlet conditions of the low pressure turbine, resulting in efficiency reduction of low pressure turbine. Highly efficient intermediate turbine duct cannot be designed without considering the effects of the rotating row of the high pressure turbine.
A typical turbine model is simulated by commercial computational fluid dynamics method. This model is used to validate the accuracy and reliability of the selected numerical method by comparing the numerical results with the experimental results. An intermediate turbine duct with eight struts has been designed initially downstream of an existing high pressure turbine. On the basis of the original design, the main purpose of this paper is to reduce the net aerodynamic load on the strut surface and thus minimize the overall duct loss. Full three-dimensional inverse method is applied to the redesign of the struts. It is revealed that the duct with new struts after inverse design has an improved performance as compared with the original one.
Numerical Analysis of an Alternate Stall in a Radial Vaned Diffuser
Authors: Benichou E. , Trébinjac I. 
Published: September 2016

The flow structure in the radial diffuser of a centrifugal compressor is analyzed from steady and unsteady Reynolds-Averaged Navier Stokes (RANS) simulations performed at one rotation speed for which two stable operating ranges separated by an unstable zone have been experimentally experienced.
Below a given mass flow rate, close to the peak efficiency point, phase-lagged single-passage simulations do not converge properly anymore. A low frequency appears in the CFD, which cannot be associated with any physical phenomenon. The computational domain is then extended, so that several passages of both the impeller and the diffuser are taken into account. At intermediate mass flow rates, an unstable operating range exists and simulations cannot converge properly either. Nevertheless, if the compressor is further throttled, another stable operating range is obtained at low mass flow rates. The flow structure in that stable operating range is rather unusual: an internal periodicity emerges inside the radial diffuser, involving a two-channel flow pattern.
This two-channel flow pattern is found to be stable and fixed in time. Moreover, the phase shift between two adjacent channel pairs happens to be constant. This indicates that a new space – time periodicity is established at low mass flow rates, which involves groups of two passages in the radial diffuser. It is confirmed thanks to a new phase-lagged simulation including one impeller passage and two diffuser passages which shows a good convergence and which gives the same results, both in terms of performance and flow physics.
Formation Process of Water Film and Performance Effect on Compressor Stage
Authors: Zhang Hai , Tian Xiaojiang , Pan Xiaojun , Zhou Jie , Zheng Qun 
Published: September 2016

In process of wet compression, gas turbine engine will ingest a certain amount of water, which can influence the overall performance of the engine. This phenomenon is particularly significant in the cleaning process of industrial gas turbine and water injection of aero-engine. When the quantity of water ingestion is quite large, the performance of gas turbine will appear deterioration and may lead to flameout, power reduce or even shutdown of the engine, causing accidents. Water droplets will be accumulated on the blade surface where water films could be formed on pressure surface in the wet compression process. The effects of water film on gas turbine engines are aerodynamic, thermodynamic and mechanical. The above-mentioned effects occur simultaneously and be affected by each other. Considering the above effects and the fact that they are time dependent, there are few gas turbine performance researches, which take into account the water film phenomenon.
This study is a new research of investigating theoretically the water film effects on a gas turbine performance. It focuses on the aerodynamic and thermodynamic effects of the phenomenon on the compressor stage. The computation of water film thickness, which frequently be formed on the surface of compressor blade, its movement and extra torque demand, are provided by a simulation model of the code. Considering the change in blade’s profile and the thickness feature of the water film, the compressor stage’s performance deterioration is analyzed. In addition to this, movement and the formation of the water film on a compressor stage are simulated and analyzed by using unsteady numerical methods under different water injecting conditions in this paper. The movement characteristics of water droplets in compressor passage are investigated to understand the flow mechanisms responsible for water film formation process. The forming and the tearing process of water film on blade surface are analyzed at different injection conditions. For simulating the real situation, The maximum quantity of injected water can reach 12%.
The results indicate that continuity and region of the water film on the blade surface will be developed with the increment of droplet size and injection rate. It is also found that the flow losses near blade surface increases with the tearing process of water film due to the increment of surface roughness.
A 3D Coupled Approach for the Thermal Design of Aero-Engine Combustor Liners
Authors: Mazzei L. , Andreini A. , Facchini B. , Bellocci L. 
Published: September 2016

The adoption of lean burn combustion to limit NOx emissions of modern aero-engines imposes a drastic reduction of air dedicated to cooling combustor dome and liners. In the latest years many aero-engine manufacturers are hence implementing effusion cooling, which provides uniform protection on the hot side of the liner and significant heat removal within the perforation. With an industrial perspective, the development of such components is usually carried out with different strategies depending on the level of accuracy required in the design phase involved (i.e preliminary or detailed). In the collaboration between GE Avio and University of Florence, the preliminary design of these devices is carried out with Therm1D, an in-house thermal flow-network solver based on the 1D correlative approach proposed by Lefebvre. This strategy, however, is not capable of taking into account the complexity of the three-dimensional nature of the flow field and the interaction between swirling flow and liner cooling, making necessary the use of Computational Fluid Dynamics (CFD) in the most advanced phases of the design process. Nevertheless, notwithstanding the increasing popularity of CFD, even a RANS simulation of a single sector of an annular combustor still presents a challenge, when the cooling system is taken into account. This issue becomes more critical in case of modern effusion cooled combustors, which may contain thousands of holes for each sector.
With the aim of of increasing the fidelity of the prediction, keeping in mind the industrial needs for limited computational efforts, a new tool has been developed: Therm3D. This approach involves the CFD simulation of the combustor flametube by modelling effusion cooling with point mass sources, whereas the fluid dynamic prediction of the remaining part is fulfilled exploiting the equivalent flow-network solver implemented in Therm1D, which provides the estimation of flow split and cold side heat loads. The solution is coupled with two separate calculations aimed at solving flame radiation and heat conduction within the metal. This paper describes the main findings of the application of Therm3D to a lean annular combustor. The results obtained have been compared to experimental data and the above mentioned numerical tools employed during the design process.
Fast Settling Millimetre-Scale Five-Hole Probes
Authors: Grimshaw S. D. , Taylor J. V. 
Published: September 2016

This paper investigates how the external geometry of a five-hole probe affects its accuracy and how internal geometry affects its settling time. An analytical model, which predicts settling time, is used to design an accurate, fast-settling, millimetre-scale probe. The paper has three components:
First, results are presented from a series of area traverses performed with five-hole probes which range in head diameter from 0.99 to 2.67 mm. It is found that the smallest probe gives the greatest accuracy when traversing the shear layers in blade wakes. However, it takes 3.4 times longer to complete this traverse than compared to the largest probe. This is because traverses with small probes require more time to allow the pressure readings to settle between each traverse position.
Second, an analytical model is developed which predicts settling time based upon the internal geometry of the probe. The approach adopted is capable of modeling any number of connected tubes with different lengths and diameters. It is validated against experimental measurements and is shown to give good agreement. This model can be used to ensure that probes are designed with acceptable settling times.
Finally, the analytical model is used to design an optimised five hole probe. Use of the model highlights two important results which are required to reduce settling time: First, the length of the smallest diameter tubes, i.e. the ones in the probe head, should be minimised. Second, the volume of tubing downstream of the head should be minimised. Applying these principles to a new probe design cuts the total traverse time by 71%, whilemaintaining the highest value of accuracy.
Spray Characteristics Measured in a New FLOX® Based Low Emission Combustor for Liquid Fuels Using Laser and Optical Diagnostics
Authors: Gounder James D. , Zizin Anton , Lammel Oliver , Aigner Manfred 
Published: September 2016

A liquid fuel combustor based on the FLOX® burner concept has been developed for application in a Micro Gas Turbine (MGT) Range Extender (REX) for next generation cars. The characterization of this combustor was performed at the High Pressure Optical Test rig (HIPOT) at DLR Stuttgart. The operability limits of the burner were mapped out for full load conditions at 3.5 bars by varying global lambda (λG) from 1.25–2.00 and bulk jet velocity (vBulk) from 80–140 m/s. Exhaust gas measurements show NOx and CO levels below 5 and 10 ppm respectively (corrected for reference 15% O2) at λG = 1.89.
Optical and laser diagnostic measurement techniques have been employed to characterize the spray flames. The flames at stable burner operation points (BOPs) show a predominantly jet like flame shape irrespective of λG and vBulk. Droplets in the size range 2–40 μm have been measured close to the nozzle exit plane. Velocities conditioned on the droplet size show large droplets d > 15 μm transitioning from negative slip velocity at the exit plane to positive slip velocity at downstream location. The positive slip velocities and slow evaporation of large droplets lead to droplets travelling further into the combustion chamber and hence resulting in long flames. A comprehensive data set for the spray characteristic of the new liquid FLOX® burner is made available.
Full-Scale Turbofan Demonstration of a Deployable Engine Air-Brake for Drag Management Applications
Authors: Shah Parthiv N. , Pfeiffer Gordon , Davis Rory , Hartley Thomas , Spakovszky Zoltán 
Published: September 2016

This paper presents the design and full-scale ground-test demonstration of an engine air-brake (EAB) nozzle that uses a deployable swirl vane mechanism to switch the operation of a turbofan’s exhaust stream from thrust generation to drag generation during the approach and/or descent phase of flight. The EAB generates a swirling outflow from the turbofan exhaust nozzle, allowing an aircraft to generate equivalent drag in the form of thrust reduction at a fixed fan rotor speed. The drag generated by the swirling exhaust flow is sustained by the strong radial pressure gradient created by the EAB swirl vanes. Such drag-on-demand is an enabler to operational benefits such as slower, steeper, and/or aeroacoustically cleaner flight on approach, addressing the aviation community’s need for active and passive control of aeroacoustic noise sources and access to confined airports.
Using NASA’s Technology Readiness Level (TRL) definitions, the EAB technology has been matured to a level of 6, i.e., a fully functional prototype. The TRL-maturation effort involved design, fabrication, assembly, and ground-testing of the EAB’s deployable mechanism on a full-scale, mixed-exhaust, medium-bypass-ratio business jet engine (Williams International FJ44-4A) operating at the upper end of typical approach throttle settings. The final prototype design satisfied a set of critical technology demonstration requirements that included (1) aerodynamic equivalent drag production equal to 15% of nominal gross thrust in a high-powered approach throttle setting (called dirty approach), (2) excess nozzle flow capacity and fuel burn reduction in the fully deployed configuration, (3) acceptable engine operability during dynamic deployment and stowing, (4) deployment time of 3–5 seconds, (5) stowing time under 0.5 second, and (6) packaging of the mechanism within a notional engine cowl. For a typical twin-jet aircraft application, a constant-speed, steep approach analysis suggests that the EAB drag could be used without additional external airframe drag to increase the conventional glideslope from 3 to 4.3 degrees, with about 3 dB noise reduction at a fixed observer location.
Experimental Investigation of Variable Geometry Turbine Annular Cascade for Marine Gas Turbines
Authors: Niu Xiying , Liang Chen , Jing Xuemei , Wei Jia , Zhu Kaidi 
Published: September 2016

Gas turbines are widely used as the marine main power system with its higher power density, react quickly, such as LM2500 and MT30. However, it works under design conditions only during running times of 3% to 10%, and it works under part load during most of the time, leading to low efficiency, and it could not achieve full speed or braking at an instant if sudden emergencies happen. Variable geometry turbines can improve this condition by variable angle nozzle (VAN) technology. And, it could enhance engine braking ability, reduce the fuel consumption under part load, improve the aerodynamic performance of engines, enhance accelerating ability of engines, and implement stalling protection to the power turbine.
However, the VAN adjustment needs complicated regulating systems, which makes it difficult to turbine structural design, and leads to increased weight. Besides, there is a performance penalty associated with the vane-end part radial clearance required for the movement of variable vanes.
In order to increase the part load efficiency of an intercooled recuperated gas turbine, the power turbine is converted from fixed to variable geometry. And, in order to reduce the losses caused by the radial clearance both of vane ends while vane turning, spherical ends are introduced to keep the clearance constant at all turning angles, and the baseline clearance is 0.77% of blade span.
In order to determine the effects of VAN on aerodynamic performance of a variable vane, experimental investigations with a variable geometry turbine annular sector cascade have been conducted under five different turning angles (−6°, −3°, 0°, +5° and +10°) and three Mach numbers (0.3Ma, 0.5Ma and 0.6Ma). The parameter distributions were measured at cascade downstream by a five-hole probe and three-axis auto-traversing system, including outlet flow angle, total pressure loss coefficient, energy loss coefficient.
The sector measurement results show that, as the vane turning angle is changed from closed to open, the outlet flow angle are increased under all three test Mach number conditions, which affects the flow mismatching between variable vane and downstream row. And, the total pressure losses is increased with the turning angle changed from design to closed or open, and the total pressure loss increases much more when the vane is closed than when it is open. In addition, vane-end clearances have significantly effects on the flow field. Especially on the hub, the leakage loss is higher, that may be due to the adverse effect of intermediate turbine ducts. Detailed results about these are presented and discussed in the paper.
Reduction of Turbine Blade Unsteady Forces by Shape Modification of Vanes for Industrial Gas Turbines
Authors: Niu Xiying , Wang Lin , Li Dongming , Du Qiuli 
Published: September 2016

The flow in turbomachinery is inherently three-dimensional unsteady and turbulent. Unsteady factors due to the viscous wakes and potential effects of blades, affect the blade surface pressure distribution, which leads to blade vibrations by periodic pulsating flow forces. While the aerodynamic excitation frequency equals to the natural frequency of blades, critical blade vibrations are excited, which could lead to a reduction of lifetime or even a destruction of rotor blades.
Although a large number of investigations about turbine unsteady flow fields and unsteady excitation forces have been carried out, these investigations do not focus on how to reduce the turbine blade unsteady forces. In this article, the turbine stators were redesigned by forward-swept and positive-leaned vanes, in order to reduce the blade excitation forces. And, the swept angles and leaned angles are determined respectively, that is, the vane is forward swept 3 degree, and positive leaned 13 degree, with the purpose of ensuring that while the hub exit flow of a vane aligns with the leading edge of one downstream rotor blade, the tip exit flow could align with the leading edge of another rotor blade that is adjacent to the above rotor blade pressure side. Comparative investigations into the turbine unsteady flow fields between the redesigned and original cases are performed by using a three-dimensional Navier-Stokes viscous solver. Emphasis is placed on how the vane modification reduces the turbine blade unsteady forces.
The results indicate that the static pressure fluctuation at vane trailing edge region is reduced compared to the baseline, and high loss region inside vane passages is also reduced. Besides, blade-to-blade entropy distribution of the turbine stage shows that, the vane wake width is reduced by its shape modification, especially at 5% and 95% span. Due to these, the temporal-spatial profile of rotor blade static pressure coefficient shows that, the unsteady fluctuation of rotor blade becomes weak obviously for the redesigned case. Overall, compared to the baseline, on the condition that the mass flowrate and expansion ratio are nearly the same, the torque fluctuation is reduced from 12.78% to 6.92%, and the axial force fluctuation is reduced from 8.82% to 6.51%, which not only is good for reducing the stream exciting force, but also can make the output power more stable. And, the turbine stage time-averaged efficiency increases slightly by 0.18%. Detailed results about static pressure coefficient distributions, entropy distributions, torque and axial forces are presented and discussed in the paper.