On the Use of an Inflatable Rubber Lip to Improve the Reverse Thrust Flow Field in a Variable Pitch Fan

2021 ◽  
Author(s):  
David John Rajendran ◽  
Vassilios Pachidis
Keyword(s):  
Flow Field ◽  
Reverse Flow ◽  
Separated Flow ◽  
Geometric Feature ◽  
Engine Operation ◽  
Design Modification ◽  
Variable Pitch ◽  
Aircraft Landing ◽  
Turn Radius ◽  
Reverse Thrust

Abstract The installed Variable Pitch Fan (VPF) reverse thrust flow field is obtained from the flow solution of an integrated airframe-engine-VPF research model for the complete reverser engagement regime during the aircraft landing run. The reverse thrust flow field indicates that the reverse flow out of the nacelle inlet is washed downstream by the freestream. Consequently, reverse flow enters the engine through the bypass nozzle from a 180° turn of the washed-down stream. This results in a region of separated flow at the nozzle lip that acts as a blockage to the reverse flow entry into the engine. To mitigate the blockage issue, a smooth guidance of the reverse flow into the engine can be achieved by using an inflatable rubber lip that would define a bell-mouth like geometric feature with a round radius at the nacelle exit. In nominal engine operation, the rubber lip would be stowed flush within the contours of the nacelle surface. The design space of the rubber lip is studied by considering different rounding radii and locations of the turn radius with respect to the nacelle trailing edge. It is observed that a rounding radius of 0.1x nacelle length is sufficient to reduce the blockage and increase the ingested reverse flow by 47% to 18% in the 140 to 40 knots landing speed range. The inflatable rubber lip represents a design modification that can improve VPF reverse thrust operation, in cases where an augmentation of reverse thrust capability is desired

On the Use of an Inflatable Rubber Lip to Improve the Reverse Thrust Flow Field in a Variable Pitch Fan

2021 ◽  
Author(s):  
David John Rajendran ◽  
Vassilios Pachidis
Keyword(s):  
Flow Field ◽  
Reverse Flow ◽  
Separated Flow ◽  
Design Parameters ◽  
Geometric Feature ◽  
Engine Operation ◽  
Design Modification ◽  
Variable Pitch ◽  
Path Representation ◽  
Reverse Thrust

Abstract The installed Variable Pitch Fan (VPF) reverse thrust flow field is obtained from the flow solution of an integrated airframe-engine research model for the complete reverser engagement regime during the aircraft landing run from 140 knots to 40 knots. The model includes a twin-engine airframe, complete flow path representation of a future 40000 lbf high bypass ratio geared turbofan engine, and a bespoke reverse flow-capable VPF design. The reverse thrust flow field, at all speeds, indicates that the reverse flow out of the nacelle inlet is washed downstream by the freestream towards the engine exit regions. Consequently, reverse flow enters the engine through the bypass nozzle from a 180° turn of the washed-down stream. This results in a region of circumferentially varying separated flow at the nozzle lip that acts as a blockage to the reverse flow entry into the engine. To mitigate the blockage issue, a smooth guidance of the reverse flow into the engine to avoid separation can be achieved by using an inflatable rubber lip that would define a bell-mouth like geometric feature with a round radius at the nacelle exit region. In nominal engine operation, the rubber lip would be stowed flush within the contours of the optimized nacelle surface. The design space of the rubber lip is studied by considering different rounding radii and locations of the turn radius with respect to the nacelle trailing edge. The choices of the design parameters are chosen by considering the nacelle edge thickness, inflation air volume requirement, weight, and thickness of support structures. The effect of these designs on the reverse thrust flow field is studied by incorporating the designs into the integrated model, with realistic installation related restrictions. It is observed that a rounding radius of 0.1x nacelle length is sufficient to reduce the blockage and increase the ingested reverse flow by 47% to 18% in the 140 to 40 knots landing speed range. The inflatable rubber lip represents a design modification that can aid in the improvement of VPF reverse thrust operation, in cases where an augmentation of reverse thrust capability over the baseline is desired.

scholarly journals Flow Distortion Into the Core Engine for an Installed Variable Pitch Fan in Reverse Thrust Mode

2020 ◽  
Author(s):  
David John Rajendran ◽  
Vassilios Pachidis
Keyword(s):  
Flow Field ◽  
Total Pressure ◽  
Reverse Flow ◽  
Ground Plane ◽  
Operating Conditions ◽  
Flow Distortion ◽  
Engine Operation ◽  
Variable Pitch ◽  
The Core ◽  
Reverse Thrust

Abstract The flow distortion at core engine entry for a Variable Pitch Fan (VPF) in reverse thrust mode is described from a realistic flow field obtained using an integrated airframe-engine model. The model includes the VPF, core entry splitter, complete bypass nozzle flow path wrapped in a nacelle and installed to an airframe in landing configuration through a pylon. A moving ground plane to mimic the rolling runway is included. 3D RANS solutions are generated at two combinations of VPF stagger angle and rotational speed settings for the entire aircraft landing run from 140 to 20 knots. The internal reverse thrust flow field is characterized by bypass nozzle lip separation, pylon wake and recirculation of flow turned back from the VPF. A portion of the reverse stream flow turns 180° with separation at the splitter leading edge to feed the core engine. The core engine feed flow exhibits circumferential and radial non-uniformities that depend on the reverse flow development at different landing speeds. The temporal dependence of the distorted flow features is also explored by an URANS analysis. Total pressure and swirl angle distortion descriptors, as defined by the Society of Automotive Engineers (SAE) S-16 committee, and, total pressure loss into the core engine are described for the core feed flow at different operating conditions and landing speeds. It is observed that the radial intensity of total pressure distortion is critical to core engine operation, while the circumferential intensity is within acceptable limits. Therefore, the baseline sharp splitter edge is replaced by two larger rounded splitter edges of radii, ∼0.1x and ∼0.2x times the core duct height. This was found to reduce the radial intensity of total pressure distortion to acceptable levels. The description of the installed core feed flow distortion, as described in this study, is necessary to ascertain stable core engine operation, which powers the VPF in reverse thrust mode.

scholarly journals Flow Distortion into the Core Engine for an Installed Variable Pitch Fan in Reverse Thrust Mode

2021 ◽  
pp. 1-30
Author(s):  
David John Rajendran ◽  
Vassilios Pachidis
Keyword(s):  
Total Pressure ◽  
Reverse Flow ◽  
Flow Distortion ◽  
Engine Operation ◽  
Variable Pitch ◽  
The Core ◽  
Aircraft Landing ◽  
Swirl Angle ◽  
Flow Features ◽  
Reverse Thrust

Abstract The flow distortion at core engine entry for a Variable Pitch Fan (VPF) in reverse thrust mode is described from a realistic flowfield obtained using an integrated airframe-engine-VPF research model. 3D RANS solutions are generated for the complete aircraft landing run from 140 to 20 knots at different VPF settings. The internal reverse thrust flowfield is characterized by nozzle lip separation, pylon wake and recirculation of flow turned back from the VPF. A portion of the reverse flow turns 180° with separation at the splitter edge to feed the core engine. The core feed flow exhibits circumferential and radial non-uniformities that depend on the reverse flow development at different landing speeds. The temporal dependence of the distorted flow features is also explored by an URANS analysis. Total pressure and swirl angle distortion descriptors, and total pressure loss are described for the core feed flow at different VPF settings and landing speeds. It is observed that the radial intensity of total pressure distortion is critical to core engine operation, while the circumferential intensity is within acceptable limits. Therefore, the baseline sharp splitter edge is replaced by two larger rounded splitter edges of radii, ∼0.1x and ∼0.2x times the core duct height. This was found to reduce the radial intensity of total pressure distortion to acceptable levels. The description of the installed core feed flow distortion, as in this study, is necessary to ascertain stable core engine operation, which powers the VPF in reverse thrust mode.

scholarly journals Fan Flow Field in an Installed Variable Pitch Fan Operating in Reverse Thrust for a Range of Aircraft Landing Speeds

2019 ◽  
Author(s):  
David John Rajendran ◽  
Vassilios Pachidis
Keyword(s):  
Flow Field ◽  
Ground Plane ◽  
Variable Pitch ◽  
Stream Flows ◽  
Engine Model ◽  
Aircraft Landing ◽  
Exhaust Duct ◽  
Actual Flow ◽  
Reverse Thrust ◽  
Stagger Angle

Abstract The installed flow field for a Variable Pitch Fan (VPF) operating in reverse thrust for the complete aircraft landing run is described in this paper. To do this, a VPF design to generate reverse thrust by reversing airflow direction is developed for a representative 40000 lbf modern high bypass ratio engine. Thereafter, to represent the actual flow conditions that the VPF would face, an engine model that includes the nacelle, core inlet splitter, outlet guide vanes, bypass nozzle, core exhaust duct, aft-body plug and core nozzle is designed. The engine model with the VPF is attached to a representative airframe in landing configuration to include the effects of installation. A rolling ground plane that mimics the runway during the landing run is also included to complete the model definition. 3D RANS solutions are carried out for two different VPF stagger angle settings and rotational speeds to obtain the fan flow field. The dynamic installed VPF flow field is characterized by the interaction of the free stream and the reverse stream flows. The two streams meet in a shear layer in the fan passages and get deflected radially outwards before turning back onto themselves. The flow field changes with stagger setting, fan rotational speed and the aircraft landing speed because of the consequent changes in the momentum of the two streams. The description of the installed VPF flow field as generated in this study is necessary to: a) qualify VPF designs that are typically designed by considering only the uninstalled static flow field b) choose the VPF operating setting for different stages of the aircraft landing run.

scholarly journals Fan Flow Field in an Installed Variable Pitch Fan Operating in Reverse Thrust for a Range of Aircraft Landing Speeds

2019 ◽  
Vol 141 (10) ◽  
Author(s):  
David John Rajendran ◽  
Vassilios Pachidis
Keyword(s):  
Flow Field ◽  
Ground Plane ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Variable Pitch ◽  
Stream Flows ◽  
Engine Model ◽  
Aircraft Landing ◽  
Actual Flow ◽  
Reverse Thrust

Abstract The installed flow field for a variable pitch fan (VPF) operating in reverse thrust for the complete aircraft landing run is described in this paper. To do this, a VPF design to generate reverse thrust by reversing airflow direction is developed for a representative 40,000 lbf modern high bypass ratio engine. Thereafter, to represent the actual flow conditions that the VPF would face, an engine model that includes the nacelle, core inlet splitter, outlet guide vanes, bypass nozzle, core exhaust duct, aft-body plug, and core nozzle is designed. The engine model with the VPF is attached to a representative airframe in landing configuration to include the effects of installation. A rolling ground plane that mimics the runway during the landing run is also included to complete the model definition. Three-dimensional (3D) Reynolds-averaged Navier–Stokes (RANS) solutions are carried out for two different VPF stagger angle settings and rotational speeds to obtain the fan flow field. The dynamic installed VPF flow field is characterized by the interaction of the freestream and the reverse stream flows. The two streams meet in a shear layer in the fan passages and get deflected radially outward before turning back onto themselves. The flow field changes with stagger setting, fan rotational speed, and the aircraft landing speed because of the consequent changes in the momentum of the two streams. The description of the installed VPF flow field as generated in this study is necessary to (a) qualify VPF designs that are typically designed by considering only the uninstalled static flow field and (b) choose the VPF operating setting for different stages of the aircraft landing run.

scholarly journals A New Method to Determine the Impact of Individual Field Quantities on Cycle-to-Cycle Variations in a Spark-Ignited Gas Engine

Energies ◽  
2021 ◽  
Vol 14 (14) ◽  
pp. 4136
Author(s):  
Clemens Gößnitzer ◽  
Shawn Givler
Keyword(s):  
Kinetic Energy ◽  
Flow Field ◽  
Turbulent Kinetic Energy ◽  
Negative Impact ◽  
Operating Conditions ◽  
Engine Operation ◽  
Gas Engine ◽  
Cycle Simulation ◽  
Simulation Results ◽  
The Impact

Cycle-to-cycle variations (CCV) in spark-ignited (SI) engines impose performance limitations and in the extreme limit can lead to very strong, potentially damaging cycles. Thus, CCV force sub-optimal engine operating conditions. A deeper understanding of CCV is key to enabling control strategies, improving engine design and reducing the negative impact of CCV on engine operation. This paper presents a new simulation strategy which allows investigation of the impact of individual physical quantities (e.g., flow field or turbulence quantities) on CCV separately. As a first step, multi-cycle unsteady Reynolds-averaged Navier–Stokes (uRANS) computational fluid dynamics (CFD) simulations of a spark-ignited natural gas engine are performed. For each cycle, simulation results just prior to each spark timing are taken. Next, simulation results from different cycles are combined: one quantity, e.g., the flow field, is extracted from a snapshot of one given cycle, and all other quantities are taken from a snapshot from a different cycle. Such a combination yields a new snapshot. With the combined snapshot, the simulation is continued until the end of combustion. The results obtained with combined snapshots show that the velocity field seems to have the highest impact on CCV. Turbulence intensity, quantified by the turbulent kinetic energy and turbulent kinetic energy dissipation rate, has a similar value for all snapshots. Thus, their impact on CCV is small compared to the flow field. This novel methodology is very flexible and allows investigation of the sources of CCV which have been difficult to investigate in the past.

The Effect of Four Part Gap Geometry Configurations for Variable Stator Vanes in a Compressor Cascade

2012 ◽  
Author(s):  
Marcel Gottschall ◽  
Konrad Vogeler ◽  
Ronald Mailach
Keyword(s):  
Flow Field ◽  
Separated Flow ◽  
Reynolds Numbers ◽  
Compressor Cascade ◽  
Field Development ◽  
Operating Range ◽  
Rear Part ◽  
Axis Position ◽  
Stagger Angle ◽  
Variable Stator

The article describes numerical investigations on the influence of four different endwall clearance topologies for variable stator vanes to secondary flow field development and the performance of high pressure compressors. The aim of this work is to quantify the characteristics of different clearance configurations depending on the penny-axis position and the penny diameter for a typical operating range. All clearance configurations were implemented to a linear cascade of modern stator profiles. The analysis was introduced using a relative clearance size of 1.3% chord at three stagger angles and two characteristic Reynolds numbers to model the operating range on aircraft engines. 3D numerical calculations were carried out to gain information about the flow field inside the cascade. They were compared with measurements of a 5-hole-probe as well as pressure tappings on the airfoil and the endwall. The CFD shows the clearance characteristics in good agreement with the measurements for the lower and the nominal stagger angle. Small gaps in the rear part of the vane have a beneficial effect on the flow field. In contrast, a clearance in the higher loaded front part of the vane always resulted in increased losses. Otherwise, the significant enhanced performance of a rear part gap, which was measured at the higher stagger angle, was not reflected by the CFD. The reduced mixing losses and the higher averaged flow turning even compared to a configuration without a clearance are not verified with the calculations. Large flow separations at the high stagger angle result in a two to four times higher underturning of the CFD in comparison to the experiments. The clearance effects to the characteristic radial loss distribution up to 40 % bladeheight also deviate from the measurements due to heavy mixing of clearance and reversed separated flow.

An Investigation of Flow Fields Over Multi-Element Aerofoils

2001 ◽  
Vol 124 (1) ◽  
pp. 154-165 ◽  
Author(s):  
S. R. Maddah ◽  
H. H. Bruun
Keyword(s):  
Flow Field ◽  
Computational Study ◽  
Separated Flow ◽  
Mean Flow ◽  
Model Configuration ◽  
Attached Flow ◽  
The Mean ◽  
Flap Deflection ◽  
Comparative Results ◽  
Lift And Drag

This paper presents results obtained from a combined experimental and computational study of the flow field over a multi-element aerofoil with and without an advanced slat. Detailed measurements of the mean flow and turbulent quantities over a multi-element aerofoil model in a wind tunnel have been carried out using stationary and flying hot-wire (FHW) probes. The model configuration which spans the test section 600mm×600mm, is made of three parts: 1) an advanced (heel-less) slat, 2) a NACA 4412 main aerofoil and 3) a NACA 4415 flap. The chord lengths of the elements were 38, 250 and 83 mm, respectively. The results were obtained at a chord Reynolds number of 3×105 and a free Mach number of less than 0.1. The variations in the flow field are explained with reference to three distinct flow field regimes: attached flow, intermittent separated flow, and separated flow. Initial comparative results are presented for the single main aerofoil and the main aerofoil with a nondeflected flap at angles of attacks of 5, 10, and 15 deg. This is followed by the results for the three-element aerofoil with emphasis on the slat performance at angles of attack α=10, 15, 20, and 25 deg. Results are discussed both for a nondeflected flap δf=0deg and a deflected flap δf=25deg. The measurements presented are combined with other related aerofoil measurements to explain the main interaction of the slat/main aerofoil and main aerofoil/flap both for nondeflected and deflected flap conditions. These results are linked to numerically calculated variations in lift and drag coefficients with angle of attack and flap deflection angle.

scholarly journals Reverse Thrust Aerodynamics of Variable Pitch Fans

2018 ◽  
Author(s):  
Tim S. Williams ◽  
Cesare A. Hall
Keyword(s):  
High Speed ◽  
Pressure Ratio ◽  
Test Case ◽  
Suction Surface ◽  
Flow Angle ◽  
Variable Pitch ◽  
Open Rotor ◽  
Weight Penalty ◽  
Reverse Thrust ◽  
Rotor Loss

Variable pitch fans are of interest for future low pressure ratio fan systems since they provide improved operability relative to fixed pitch fans. If they can also be re-pitched such that they generate sufficient reverse thrust they could eliminate the engine drag and weight penalty associated with bypass duct thrust reversers. This paper sets out to understand the details of the 3D fan stage flow field in reverse thrust operation. The study uses the Advanced Ducted Propulsor variable pitch fan test case, which has a design fan pressure ratio of 1.29. Comparison with spanwise probe measurements show that the computational approach is valid for examining the variation of loss and work in the rotor in forward thrust. The method is then extended to a reverse thrust configuration using an extended domain and appropriate boundary conditions. Computations, run at two rotor stagger settings, show that the spanwise variation in relative flow angle onto the rotor aligns poorly to the rotor inlet metal angle. This leads to two dominant rotor loss sources: one at the tip associated with positive incidence, and the second caused by negative incidence at lower span fractions. The second loss is reduced by opening the rotor stagger setting, and the first increases with rotor suction surface Mach number. The higher mass flow at more open rotor settings provide higher gross thrust, up to 49% of the forward take-off value, but is limited by the increased loss at high speed.

scholarly journals Dynamics of flow in a branching channel

2021 ◽  
Vol 22 ◽  
pp. 25
Author(s):  
Václav Uruba ◽  
Pavel Procházka ◽  
Vladislav Skála
Keyword(s):  
Flow Field ◽  
Cross Section ◽  
Rectangular Cross Section ◽  
Separated Flow ◽  
Heat Transfer Process ◽  
Dynamical Behaviour ◽  
Main Channel ◽  
Flow Dynamics ◽  
Mean Flow ◽  
Piv Technique

Flow in a branched channel is studied experimentally using the PIV technique. The presented study is concentrated on clarifying the dynamical behaviour in individual branches. The 11 branches issuing from the main channel perpendicularly, all channels are of rectangular cross-section. First, the time-mean flow-field is shown, then the flow dynamics is investigated using the OPD method. Flow patterns and frequencies are evaluated in three selected branches. The separated flow in branches exhibits highly dynamical behaviour, which differs substantially in the branches close to the inflow, in the main channel middle and close to its end. The typical topologies and frequencies of the detected quasi-periodical structures in the channel braches are shown in the study. Mostly, the flow-fields are populated by trains of vortices with alternating orientation and saddle-like structures. The flow-field close to the channel walls affects heat transfer process between the wall and fluid.

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