The Influence of Fan Root Flow on the Aerodynamics of a Low-Pressure Compressor Transition Duct

2019 ◽  
Author(s):  
A. D. Walker ◽  
I. Mariah ◽  
D. Tsakmakidou ◽  
H. Vadhvana ◽  
C. Hall
Keyword(s):  
Gas Turbines ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Profile ◽  
Low Pressure ◽  
Secondary Flows ◽  
Test Facility ◽  
Inlet Flow ◽  
Flat Profile ◽  
Bypass Ratio

Abstract To reduce fuel-burn and CO2 emissions from aero gas turbines there is a drive towards very-high bypass ratio and smaller ultra-high-pressure ratio core engine technologies. However, this makes the design of the ducts connecting various compressor spools more challenging as the higher required radius change increases their aerodynamic loading. This is exacerbated for the duct which feeds the engine core as it must accept the relatively low-quality flow produced by the fan root. This is characterised by a hub-low pressure profile and large secondary flow structures which will inevitably increase loss and the likelihood of flow separation. Additionally, the desire for shorter, lighter nacelles means that the engine intake may be unable provide a uniform inlet flow to the fan when the aircraft is at an angle of attack or subject to cross winds. Any inlet distortion this generates can also further degrade the quality of the flow entering the core of the engine. This paper uses a combination of experiments and CFD to examine the effects of the inlet flow on the aerodynamics of an engine section splitter and transition duct designed to feed the low-pressure spool of a high bypass ratio turbofan. A fully annular test facility incorporating a 1½ stage axial compressor was used to compare the system performance of a rotor that produced a nominally flat profile with one that had a notably hub deficient flow. A RANS CFD model, employing a mixing plane between the rotor and Engine Section Stator (ESS) and a Reynolds Stress turbulence model, was then validated and used to further investigate the effects of increased inlet boundary layer thickness and bulk swirl distortion at rotor inlet. Overall, changes to the inlet condition were seen to have a surprisingly small effect on the flow at duct exit — i.e. the flow presented to the downstream compressor. Changes to the inlet did, however, generate increased secondary flows and degrade the performance of the ESS. This resulted in notably increased total pressure loss; in excess of 12% for the hub-low inlet and in excess of 30% at high inlet swirl where the flow in the ESS separated. However, the increased ESS wake structures, and the enhanced mixing, delayed separation in the duct suggesting that, overall the design was reasonably robust, albeit with a significant penalty in system loss.

scholarly journals The Influence of Fan Root Flow on the Aerodynamic of a Low-Pressure Compressor Transition Duct

2019 ◽  
Vol 142 (1) ◽  
Author(s):  
A. Duncan Walker ◽  
Ian Mariah ◽  
Dimitra Tsakmakidou ◽  
Hiren Vadhvana ◽  
Chris Hall
Keyword(s):  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Profile ◽  
Low Pressure ◽  
Secondary Flows ◽  
Test Facility ◽  
Navier Stokes ◽  
High Pressure Ratio ◽  
Fuel Burn ◽  
Bypass Ratio

Abstract To reduce fuel-burn and emissions, there is a drive toward higher bypass ratio and smaller high-pressure ratio core engines. This makes the design of the ducts connecting compressor spools more challenging as the higher radius change increases aerodynamic loading. This is exacerbated at inlet to the engine core by fan root flow which is characterized by a hub-low-pressure profile and large secondary flow structures. Additionally, shorter, lighter nacelles mean that the intake may not provide a uniform inlet flow when the aircraft is at an angle of attack or subject to cross winds. Such inlet distortion can further degrade the flow entering the engine. A combination of experiments and computational fluid dynamics (CFD) has been used to examine the effects on the aerodynamics of an engine section splitter (ESS) and transition duct designed to feed the low-pressure spool of a high bypass ratio turbofan. A test facility incorporating a 1½ stage axial compressor was used to compare system performance for a flat rotor exit profile to one with a hub deficient flow. Validated Reynolds averaged Navier–Stokes (RANS) CFD was then used to further investigate the effects of increased inlet boundary layer thickness and bulk swirl distortion at rotor inlet. These changes were seen to have a surprisingly small effect on the flow at the duct exit. However, increased secondary flows were observed which degraded the performance of the ESS and significantly increased loss. Nevertheless, the enhanced mixing delayed separation in the duct suggesting that overall the design was reasonably robust albeit with increased system loss.

A Numerical Investigation Into Secondary Flows at the Inlet to a Low-Pressure Compressor Transition Duct

2020 ◽  
Author(s):  
Dimitra Tsakmakidou ◽  
A. Duncan Walker ◽  
Chris Hall
Keyword(s):  
Boundary Layer ◽  
Layer Thickness ◽  
Gas Turbines ◽  
Boundary Layer Thickness ◽  
Pressure Ratio ◽  
Low Pressure ◽  
Secondary Flows ◽  
Test Facility ◽  
Rotor Loss ◽  
Bypass Ratio

Abstract To reduce fuel-burn and CO2 emissions from aero gas turbines there is a drive towards very-high bypass ratio and smaller ultra-high-pressure ratio core engine technologies. However, this makes the design of the ducts connecting various compressor spools more challenging as the larger required radius change increases their aerodynamic loading. This is exacerbated for the duct which feeds the engine core as it must accept the relatively low-quality flow produced by the fan root. This is characterised by a hub-low pressure profile and large secondary flow structures which will inevitably increase loss and the likelihood of flow separation. This paper presents a numerical design study aimed at modifying an existing fully annular, low-speed test facility to produce larger, more representative rotor loss cores. The CFD domain comprises a single stage axial compressor and transition duct representative of the low-pressure spool in a high bypass ratio turbofan. The rotating and stationary frames were coupled using a sliding mesh and a uRANS approach with Reynolds Stress closure for the turbulence modelling. The methodology was first validated using existing experimental data before examining a number of parameters such as inlet boundary layer thickness, inlet swirl, tip gap, blade solidity, thickness, lean. It was found that a combination of an increased inlet hub boundary layer and a reduction in solidity sufficiently increased the hub loading and encouraged the development of significantly large loss cores. Preliminary CFD results also showed that the increased rotor loss cores promoted better mixing through the outlet guide which subsequently reduced hub boundary layer thickness and secondary flows. Despite incurring a larger total pressure loss this suggested that the transition duct moved further from stall and could therefore potentially be made shorter.

Experimental Investigation of Secondary Flows and Length Reduction for a Low-Pressure Compressor Transition Duct

2021 ◽  
Author(s):  
D. Tsakmakidou ◽  
I. Mariah ◽  
A. D. Walker ◽  
C. Hall ◽  
H. Simpson
Keyword(s):  
Boundary Layer ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Low Pressure ◽  
Secondary Flows ◽  
Test Facility ◽  
Momentum Exchange ◽  
Fuel Burn ◽  
Operational Life ◽  
Length Reduction

Abstract The need to reduce fuel-burn and CO2 emissions, is pushing turbofan engines towards geared architectures with very high bypass-ratios and small ultra-high-pressure ratio core engines. However, this increases the radial offset between compressor spools and leads to a more challenging design for the compressor transition ducts. To minimise weight, these ducts must achieve the radial turning in as short a length, but this leads to strong curvature induced pressure gradients, increased aerodynamic loading and likelihood of flow separation. For the duct connecting the low-pressure fan to the engine core this is further complicated by the poor-quality flow generated at the fan hub which is characterised by low total pressure and large rotating secondary flow structures. In a previous paper the authors numerically designed modifications to an existing test facility such that the rotor would produce these large structures. The current paper presents an experimental evaluation of the new rotor design and examines the effect of the increased loss cores on the performance of a set of engine sector stators (ESS) or outlet guide vanes (OGV) and an engine representative compressor transition duct. Aerodynamic data were collected via miniature five-hole probes, for the time-averaged pressure and velocity field, and phase-locked hot-wire anemometry to capture the rotating secondary flows. Analysis of the experimental data showed that these structures promoted mixing through the ESS increasing the momentum exchange between the core and boundary layer flows. Measurements within the duct showed a continued reduction in the hub-wall boundary layer suggesting that the duct has been moved further from separation. Consequently, a more aggressive duct with 12.5% length reduction was designed and tested with the data confirming that the more aggressive duct remained fully attached. Total pressure loss data suggested a slight increase in loss over the vane row but that was offset by a reduced loss in the duct due to improved flow quality and reduced length. Overall, the 12.5% length reduction represents a significant cumulative effect in terms of reduced fuel burn and CO2 over the operational life of an engine.

Flow Characteristics of a Novel Centrifugal Compressor Design

2016 ◽  
Author(s):  
Shashank Mishra ◽  
Shaaban Abdallah ◽  
Mark Turner
Keyword(s):  
Centrifugal Compressor ◽  
Pressure Ratio ◽  
Flow Characteristics ◽  
Axial Compressor ◽  
Low Pressure ◽  
Flow Path ◽  
Optimization Techniques ◽  
Single Stage ◽  
Multi Stage ◽  
Compressor Design

Multistage axial compressor has an advantage of lower stage loading as compared to a single stage. Several stages with low pressure ratio are linked together which allows for multiplication of pressure to generate high pressure ratio in an axial compressor. Since each stage has low pressure ratio they operate at a higher efficiency and the efficiency of multi-stage axial compressor as a whole is very high. Although, single stage centrifugal compressor has higher pressure ratio compared with an axial compressor but multistage centrifugal compressors are not as efficient because the flow has to be turned from radial at outlet to axial at inlet for each stage. The present study explores the advantages of extending the axial compressor efficient flow path that consist of rotor stator stages to the centrifugal compressor stage. In this invention, two rotating rows of blades are mounted on the same impeller disk, separated by a stator blade row attached to the casing. A certain amount of turning can be achieved through a single stage centrifugal compressor before flow starts separating, thus dividing it into multiple stages would be advantageous as it would allow for more flow turning. Also the individual stage now operate with low pressure ratio and high efficiency resulting into an overall increase in pressure ratio and efficiency. The baseline is derived from the NASA low speed centrifugal compressor design which is a 55 degree backward swept impeller. Flow characteristics of the novel multistage design are compared with a single stage centrifugal compressor. The flow path of the baseline and multi-stage compressor are created using 3DBGB tool and DAKOTA is used to optimize the performance of baseline as well novel design. The optimization techniques used are Genetic algorithm followed by Numerical Gradient method. The optimization resulted into improvements in incidence and geometry which significantly improved the performance over baseline compressor design. The multistage compressor is more efficient with a higher pressure ratio compared with the base line design for the same work input and initial conditions.

Improving Steady CFD to Capture the Effects of Radial Mixing in Axial Compressors

2019 ◽  
Author(s):  
Lorenzo Cozzi ◽  
Filippo Rubechini ◽  
Andrea Arnone ◽  
Andrea Schneider ◽  
Pio Astrua
Keyword(s):  
Steady State ◽  
Gas Turbines ◽  
Turbulent Viscosity ◽  
Computational Cost ◽  
Axial Compressor ◽  
Secondary Flows ◽  
Tip Clearance ◽  
Axial Compressors ◽  
Rans Model ◽  
Aspect Ratios

Abstract The axial compressors of power-generation gas turbines have a high stage count, blades with low aspect ratios and relatively large clearances in the rear section. These features promote the development of strong secondary flows. An important outcome deriving from the convection of intense secondary flows is the enhanced span-wise transport of fluid properties mainly involving the rear stages, generally referred to as “radial mixing”. An incorrect prediction of this key phenomenon may result in inaccurate performance evaluation and could mislead the designers during the compressor design phase. As shown in a previous work, in the rear stages of an axial compressor the stream-wise vorticity associated with tip clearance flows is one of the main drivers of the overall span-wise transport phenomenon. Limiting it by circumferentially averaging the flow at row interfaces is the reason why a steady-state analysis strongly under-predicts radial mixing. To properly forecast the span-wise transport within the flow-path, an unsteady analysis should be adopted. However, due to the high blade count, this approach has a computational cost not yet suitable for industrial purposes. Currently, only the steady-state full-compressor simulation can fit in a lean industrial design chain and any model upgrade improving its radial mixing prediction would be highly beneficial for the daily design practice. To attain some progresses in RANS model, its inherent lack of convection of stream-wise vorticity must be addressed. This can be done by acting on another mixing driver, able to provide the same outcome, that is turbulent diffusion. In particular, by enhancing turbulent viscosity one can promote span-wise diffusion, thus improving the radial mixing prediction of the steady approach. In this paper, this strategy to update the RANS model and its application in simulations on a compressor of the Ansaldo Energia fleet is presented, together with the model tuning that has been performed using the results of unsteady simulations as the target.

Evaluation of Axial Compressor Characteristics Under Overspray Condition

2013 ◽  
Author(s):  
Chihiro Myoren ◽  
Yasuo Takahashi ◽  
Manabu Yagi ◽  
Takanori Shibata ◽  
Tadaharu Kishibe
Keyword(s):  
Gas Turbine ◽  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Water Mass ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Test Facility ◽  
Compressor Performance ◽  
Water Mass Flow Rate

An axial compressor was developed for an industrial gas turbine equipped with a water atomization cooling (WAC) system, which is a kind of inlet fogging technique with overspray. The compressor performance was evaluated using a 40MW-class test facility for the advanced humid air turbine system. A prediction method to estimate the effect of WAC was developed for the design of the compressor. The method was based on a streamline curvature (SLC) method implementing a droplet evaporation model. Four test runs with WAC have been conducted since February 2012. The maximum water mass flow rate was 1.2% of the inlet mass flow rate at the 4th test run, while the design value was 2.0%. The results showed that the WAC decreased the inlet and outlet temperatures compared with the DRY (no fogging) case. These decreases changed the matching point of the gas turbine, and increased the mass flow rate and the pressure ratio by 1.8% and 1.1%, respectively. Since prediction results agreed with the results of the test run qualitatively, the compressor performance improvement by WAC was confirmed both experimentally and analytically. The test run with the design water mass flow rate is going to be conducted in the near future.

scholarly journals Design and Test of a New Axial Compressor for the Nuovo Pignone Heavy Duty Gas Turbines

1996 ◽  
Author(s):  
Erio Benvenuti
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Dynamic Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Leading Edge ◽  
Meridional Flow ◽  
Heavy Duty ◽  
Pressure Transducers ◽  
Easy Integration

This axial compressor design was primarily focused to increase the power rating of the current Nuovo Pignone PGT10 Heavy-Duty gas turbine by 10%. In addition, the new 11-stage design favourably compares with the existing 17-stage compressor in terms of simplicity and cost. By seating the flowpath and blade geometry, the new aerodynamic design can be applied to gas turbines with different power ratings as well. The reduction in the stage number was achieved primarily through the meridional flow-path redesign. The resulting higher blade peripheral speeds achieve larger stage pressure ratios without increasing the aerodynamic loadings. Wide chord blades keep the overall length unchanged thus assuring easy integration with other existing components. The compressor performance map was extensively checked over the speed range required for two-shaft gas turbines. The prototype unit was installed on a special PGT10 gas turbine setup, that permitted the control of pressure ratio independently from the turbine matching requirements. The flowpath instrumentation included strain-gages, dynamic pressure transducers and stator vane leading edge aerodynamic probes to determine individual stage characteristics. The general blading vibratory behavior was proved fully satisfactory. With minor adjustments to the variable stator settings the front stage aerodynamic matching was optimized and the design performance was achieved.

scholarly journals The Development of Supersonic Axial Compressor Boost Stages for Small Gas Turbines

1969 ◽  
Author(s):  
C. H. Muller ◽  
A. Sabatiuk
Keyword(s):  
Gas Turbines ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Performance Goals ◽  
Ample Evidence ◽  
Rotor Type ◽  
Highly Loaded ◽  
Percent Efficiency

The axial supersonic compressors of the “shock-in-rotor” type are under development for application to small gas turbines. A passage flow approach and passage criteria were used to design and develop the airfoils for the highly loaded rotor and stator blading of these 4 lb/sec machines. The overall stage performance values demonstrated to date are 2.06:1 pressure ratio at 78 percent adiabatic efficiency and 2.56:1 at 74.4 percent efficiency. The loss data and static pressure rise measured for the rotors and exit stators provide ample evidence that the higher design performance goals can be met.

Numerical Validations of Secondary Flows and Loss Development Downstream of a Highly Loaded Low Pressure Turbine Outlet Guide Vane Cascade

2007 ◽  
Author(s):  
Johan Hja¨rne ◽  
Valery Chernoray ◽  
Jonas Larsson ◽  
Lennart Lo¨fdahl
Keyword(s):  
Numerical Simulations ◽  
Secondary Flow ◽  
Incompressible Flows ◽  
Guide Vane ◽  
Low Pressure ◽  
Flow Fields ◽  
Secondary Flows ◽  
Test Facility ◽  
Low Pressure Turbine ◽  
Highly Loaded

In this paper 3D numerical simulations of turbulent incompressible flows are validated against experimental data from the linear low pressure turbine/outlet guide vane (LPT/OGV) cascade at Chalmers in Sweden. The validation focuses on the secondary flow-fields and loss developments downstream of a highly loaded OGV. The numerical simulations are performed for the same inlet conditions as in the test-facility with engine-like properties in terms of Reynolds number, boundary-layer thickness and inlet flow angles with the goal to validate how accurately and reliably the secondary flow fields and losses for both on- and off-design conditions can be predicted for OGV’s. Results from three different turbulence models as implemented in FLUENT, k-ε Realizable, kω-SST and the RSM are validated against detailed measurements. From these results it can be concluded that the RSM model predicts both the secondary flow field and the losses most accurately.

The Evolution of Compressor and Turbine Bladings in Gas Turbine Design

1967 ◽  
Vol 89 (2) ◽  
pp. 199-205 ◽  
Author(s):  
C. Seippel
Keyword(s):  
Steam Turbine ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Parallel Evolution ◽  
Axial Compressor ◽  
Expansion Turbine ◽  
The Stability ◽  
Turbine Design ◽  
The City

The author, having been associated with the construction of gas turbines from the first 4000-kw unit delivered in 1939 to the city of Neuchaˆtel to the present time, gives some personal views on the evolution of the axial compressor and turbine bladings which are the key elements to the gas turbines. The axial compressor was created to supply air efficiently for the supercharged “Velox” boiler. It made the evolution to the modern gas turbine possible. The main problems encountered were related to the stability of flow. An enormous increase of volume capacity was achieved in the course of time. The increase of pressure ratio made special measures necessary to overcome instability at starting. The expansion turbine started on the basis of steam turbine practice and underwent a parallel evolution to large capacities. Its particular problems are related to the high temperatures of the gases.

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