Preliminary Fan Design for a Silent Aircraft

2006 ◽  
Vol 129 (1) ◽  
pp. 184-191 ◽  
Author(s):  
Daniel Crichton ◽  
Liping Xu ◽  
Cesare A. Hall
Keyword(s):  
Mach Number ◽  
Noise Reduction ◽  
High Efficiency ◽  
Pressure Ratio ◽  
Low Pressure ◽  
Shock Structure ◽  
Operating Point ◽  
Blade Passage ◽  
Fuel Burn ◽  
Primary Shock

Preliminary fan design for a functionally silent aircraft has been performed with noise reduction as the primary goal. For such an aircraft the fan design must, in addition to delivering low cruise fuel burn, enable low jet and fan source noise during takeoff. This requires the fan to be operating at low pressure ratio and high efficiency during takeoff and, for conditions where the relative tip Mach number onto the fan is supersonic, ensuring the primary shock structure is ingested into the blade passage. To meet these requirements, flyover and cruise flow coefficients are matched using a variable area nozzle at the same time as delivering low takeoff FPR. This places the sideline operating point near the shoulder of the characteristic and fixes the top of climb and cruise fan pressure ratios. For a 4-engine, 250pax, 4000nm silent aircraft this approach leads to a top of climb FPR of 1.45, requiring a 39% increase in nozzle area at takeoff. A fan rotor has been designed for this cycle with 20 blades, low tip loading during takeoff, and a 350m∕s top of climb tip speed.

Preliminary Fan Design for a Silent Aircraft

2006 ◽  
Author(s):  
Daniel Crichton ◽  
Liping Xu ◽  
Cesare A. Hall
Keyword(s):  
Mach Number ◽  
Noise Reduction ◽  
High Efficiency ◽  
Pressure Ratio ◽  
Low Pressure ◽  
Shock Structure ◽  
Operating Point ◽  
Blade Passage ◽  
Fuel Burn ◽  
Primary Shock

Preliminary fan design for a functionally silent aircraft has been performed with noise reduction as the primary goal. For such an aircraft the fan design must, in addition to delivering low cruise fuel burn, enable low jet and fan source noise during take-off. This requires the fan to be operating at low pressure ratio and high efficiency during take-off and, for conditions where the relative tip Mach number onto the fan is supersonic, ensuring the primary shock structure is ingested into the blade passage. To meet these requirements, flyover and cruise flow coefficients are matched using a variable area nozzle at the same time as delivering low take-off FPR. This places the sideline operating point near the shoulder of the characteristic and fixes the top of climb and cruise fan pressure ratios. For a 4-engine, 250pax, 4000nm silent aircraft this approach leads to a top of climb FPR of 1.45 requiring a 39% increase in nozzle area at take-off. A fan rotor has been designed for this cycle with 20 blades, low tip loading and a 350m/s top of climb tip speed.

Design and Analysis of a High Pitch to Chord Ratio Cascade Representative of Ducted Propfans

1991 ◽  
Author(s):  
A. Weber ◽  
W. Steinert ◽  
H. Starken
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
High Efficiency ◽  
Pressure Ratio ◽  
Flow Analysis ◽  
High Pitch ◽  
Integral Boundary ◽  
Pressure Loss Coefficient ◽  
Distribution Boundary ◽  
Very High

Efforts to reduce the specific fuel consumption of a modern aero engine focus in particular on increasing the by-pass ratio beyond the current level of around 5. One concept is the counterrotating shrouded propfan operating at low overall pressure ratio and having only very few fan blades of extremely high pitch/chord ratios. The relative inlet Mach numbers cover a range from 0.7 at the hub to 1.1 at the tip section of the first rotor. A propfan cascade was designed by taking into account two characteristic features of a propfan blade-blade section: • a very high pitch/chord ratio of s/c = 2.25 • an inlet Mach number of M1 = 0.90 which leads to transonic flow conditions inside the blade passage In the design process a profile generator and a quasi-3D Euler solver were used iteratively to optimize the profile Mach number distribution. Boundary layer behavior was checked with an integral boundary layer code. The cascade design was verified experimentally in the transonic cascade wind tunnel of DLR at Cologne. The extensive experimental results confirm the design goal of roughly 5 degree flow turning. A total pressure loss coefficient of less than 1.5% was measured at design conditions. This validates the very high efficiency level the propfan concept is calling for. A 2D Navier-Stokes flow analysis code yields good results in comparison to the experimental ones.

Development of a Supersonic Ejector for Capturing Very Low-Pressure Vent Gases and Re-Injection Into a High-Pressure Gas Stream

2006 ◽  
Author(s):  
K. K. Botros ◽  
J. Geerligs ◽  
H. Imran ◽  
W. Thompson
Keyword(s):  
Mach Number ◽  
Optimum Design ◽  
Nozzle Exit ◽  
Pressure Ratio ◽  
Supersonic Nozzle ◽  
Low Pressure ◽  
Fuel Gas ◽  
Dry Gas Seal ◽  
Supersonic Ejector ◽  
Exit Mach Number

The purpose of the ejector device is to capture the gas leakage from a dry-gas seal at low pressure, and re-inject it into the fuel gas line to the gas generator (without the use of compressors or rotating elements), hence providing a means to utilize the gas that would otherwise be vented to atmosphere. Implementation of this device will also have the benefit of reducing greenhouse gas emissions to the atmosphere. The primary challenge to achieve the above goal lies in the fact that the leakage gas pressure is in the range of 70–340 kPag, while the minimum pressure required upstream of the fuel gas regulator is in the range of 2400–3300 kPag. The device consists of a two-stage supersonic ejector. The first stage is highly supersonic (nozzle exit Mach number ≃ 2.54), while the second stage is moderately supersonic (nozzle exit Mach number ≃ 1.72). Several tests where conducted on various configurations of the two stages on natural gas in order to arrive at the optimum design and operating parameters. The optimum design gave an expansion pressure ratio (motive/suction) of the order of 14.0 and compression pressure ratio (discharge/suction) of around 8.1. These ratios would meet the requirement of the minimum suction and discharge pressure mentioned above. This paper presents the optimum configuration arrived at after several iterations of different geometries of the supersonic nozzles, particularly for the first stage ejector, and presents the performance test results of the integrated system. The results indicate that the device would meet the requirements of capturing the low pressure, low flow dry gas seal leakage and re-inject it into the fuel gas stream with an overall ejector efficiency (based on thermodynamic availability) of 80%.

Heat-Exchanged Propulsion Gas Turbines: A Candidate for Future Lower SFC and Reduced-Emission Military and Civil Aeroengines

2009 ◽  
Author(s):  
Colin F. McDonald ◽  
Colin Rodgers
Keyword(s):  
Gas Turbines ◽  
High Efficiency ◽  
Pressure Ratio ◽  
Thermodynamic Cycle ◽  
Simple Cycle ◽  
Specific Power ◽  
Operational Experience ◽  
Fuel Cost ◽  
Turbofan Engines ◽  
Fuel Burn

After seven decades of service the evolution of simple cycle propulsion gas turbines continues with emphasis now being placed on reduced fuel burn, lower emissions, and less noise. With compressor and turbine efficiencies near plateauing, and turbine inlet temperatures paced by materials and blade cooling technologies, improvements in SFC, specific power and weight for conventional engines (including small turboprop, and turboshaft engines and larger turbofans) will likely be incremental compared with the past. With retention of the simple cycle both evolutionary and revolutionary approaches are being taken by the aeroengine industry to improve performance, particularly reduced fuel burn in an era of high fuel cost. In this paper a further step is suggested, that is in concert with meeting performance, economic, and environmental goals of future aeroengines, namely the use of a more complex thermodynamic cycle involving recuperation for turboprop and turboshaft engines, and intercooling together with recuperation for higher pressure ratio turbofan engines. The idea of heat exchanged propulsion gas turbines is not new, but the many concepts identified from studies done periodically over the last 65 years, including the few engines that were static tested and one test flown, didn’t find acceptance in an era of low fuel cost and concerns about recuperator integrity and reliability. With today’s very high fuel cost there is current interest in this type of engine because of its potential for low SFC and reduced emissions. In this paper potential applications are summarized and the features of various heat exchanged aeroengine design concepts together with projected performance are presented. Included is a discussion on the various issues that must be resolved before they enter service. A postulated deployment scenario is suggested with engines initially developed to meet military aviation needs, such as recuperated turboprop and turbofan engines for extended range UAV’s, followed by a recuperated turboshaft engine for a helicopter. Operational experience and demonstrated reliability from these would pave the way for high efficiency ICR turbofan engines for military and civil aircraft service sometime after the year 2020.

Sweep Effects on Fan-Intake Aerodynamics at High Angle of Attack

2021 ◽  
Author(s):  
Ben Mohankumar ◽  
Cesare A. Hall ◽  
Mark J. Wilson
Keyword(s):  
Mach Number ◽  
Pressure Ratio ◽  
Angle Of Attack ◽  
Pressure Rise ◽  
Low Pressure ◽  
Rotating Stall ◽  
High Angle ◽  
High Angle Of Attack ◽  
Mach Numbers ◽  
High Mach

Abstract Sweep in a transonic fan is conventionally used to reduce design point losses by inclining the passage shock relative to the incoming flow. However, future low pressure ratio fans operate to lower Mach numbers meaning the role of sweep at cruise is diminished. Instead, sweep might be repurposed to improve the performance of critical high Mach number off-design conditions such as high angle of attack (AOA). In this paper, we use unsteady computational fluid dynamics to compare two transonic low pressure ratio fans, one radially stacked and one highly swept, coupled to a short intake design, at the high AOA flight condition. The AOA considered is 35°, which is sufficient to separate the intake bottom lip. The midspan of the swept fan was shifted upstream to add positive sweep to the outer span. Based on previous design experience, it was hypothesised the swept fan would reduce transonic losses when operating at high AOA. However, it was found the swept fan increased the rotor loss by 24% relative to the radial fan. Loss was increased through two key mechanisms. i) Rotor choking: flow is redistributed around the intake separation and enters the rotor midspan with high Mach numbers. Sweeping the fan upstream reduced the effective intake length, which increased the inlet relative Mach number and amplified choking losses. ii): Rotor-separation interaction (RSI): the rotor tip experiences low mass flow inside the separation, which increases the pressure rise across the casing to a point where the boundary layer separates. The swept fan diffused the casing streamtube, causing the casing separation to increase in size and persist in the passage for longer. High RSI loss indicated the swept fan was operating closer to the rotating stall point.

scholarly journals Ultrashort Nacelles for Low Fan Pressure Ratio Propulsors

2014 ◽  
Vol 137 (2) ◽  
Author(s):  
Andreas Peters ◽  
Zoltán S. Spakovszky ◽  
Wesley K. Lord ◽  
Becky Rose
Keyword(s):  
Mach Number ◽  
Pressure Ratio ◽  
Design Tool ◽  
Operating Conditions ◽  
Design Parameters ◽  
Flow Distortion ◽  
Flow Area ◽  
Propulsive Efficiency ◽  
Flow Acceleration ◽  
Fuel Burn

As the propulsor fan pressure ratio (FPR) is decreased for improved fuel burn, reduced emissions and noise, the fan diameter grows and innovative nacelle concepts with short inlets are required to reduce their weight and drag. This paper addresses the uncharted inlet and nacelle design space for low-FPR propulsors where fan and nacelle are more closely coupled than in current turbofan engines. The paper presents an integrated fan–nacelle design framework, combining a spline-based inlet design tool with a fast and reliable body-force-based approach for the fan rotor and stator blade rows to capture the inlet–fan and fan–exhaust interactions and flow distortion at the fan face. The new capability enables parametric studies of characteristic inlet and nacelle design parameters with a short turn-around time. The interaction of the rotor with a region of high streamwise Mach number at the fan face is identified as the key mechanism limiting the design of short inlets. The local increase in Mach number is due to flow acceleration along the inlet internal surface coupled with a reduction in effective flow area. For a candidate short-inlet design with length over diameter ratio L/D = 0.19, the streamwise Mach number at the fan face near the shroud increases by up to 0.16 at cruise and by up to 0.36 at off-design conditions relative to a long-inlet propulsor with L/D = 0.5. As a consequence, the rotor locally operates close to choke resulting in fan efficiency penalties of up to 1.6% at cruise and 3.9% at off-design. For inlets with L/D < 0.25, the benefit from reduced nacelle drag is offset by the reduction in fan efficiency, resulting in propulsive efficiency penalties. Based on a parametric inlet study, the recommended inlet L/D is suggested to be between 0.25 and 0.4. The performance of a candidate short inlet with L/D = 0.25 was assessed using full-annulus unsteady Reynolds-averaged Navier–Stokes (RANS) simulations at critical design and off-design operating conditions. The candidate design maintains the propulsive efficiency of the baseline case and fuel burn benefits are conjectured due to reductions in nacelle weight and drag compared to an aircraft powered by the baseline propulsor.

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.

Conversion of Decommissioned Aircraft Gas Turbines to High-Efficiency Marine Units

1985 ◽  
Author(s):  
David Gordon Wilson ◽  
P. K. Poole ◽  
L. D. Owens ◽  
J. Baglione
Keyword(s):  
High Pressure ◽  
Gas Turbines ◽  
High Efficiency ◽  
Pressure Ratio ◽  
Slight Modification ◽  
Low Pressure ◽  
Development Costs ◽  
High Pressure Compressor ◽  
Efficiency Performance ◽  
Ceramic Honeycomb

Some helicopter and other engines have been shown in studies to be amenable to conversion to a low-pressure-ratio highly regenerated cycle. Typically the high-pressure compressor and high-pressure turbine would be removed, and shaft and ducting modifications would be made to introduce high-effectiveness rotary ceramic-honeycomb regenerators. In one case examined the low-pressure turbine could be used with slight modification; in others there would have to be reblading of the turbine stages and the manufacture of new shrouds. In another case the existing combustor could be used with little modification; in others new combustors would be required. Despite the extent of the modifications, the resulting high-efficiency performance over a large part of the power range and presumably relatively low capital and development costs could make this an attractive concept.

scholarly journals Influence of Different Gases on the Design Point of an Industrial Axial Compressor and Deduced Aerodynamical Rematching Methodology

2020 ◽  
Vol 4 ◽  
pp. 238-252
Author(s):  
Henrik Hoffmann ◽  
Lukas Stuhldreier ◽  
Ruben van Rennings ◽  
Peter Jeschke
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Operating Point ◽  
Air Compressor ◽  
Isentropic Exponent ◽  
Compressor Inlet ◽  
Downstream Flow ◽  
Carbon Dioxide Co2

This paper presents a numerical investigation of the effects of compressing various gases, for example, carbon dioxide (CO2) and methane (CH4), on an eight-stage axial air compressor. Several adaptation methods are applied to achieve a similar operating point as for air. Theoretically, the operating point depends on Mach number, flow angles, Reynolds number and isentropic exponent. Numerical results show mismatch effects which arise in the parameters using a non-adapted geometry. A rematching procedure is described, including deduced speed adjustments, in order to achieve Mach number equality at compressor inlet. Only shroud modifications are performed to rematch the flow angles of the air simulation. Although Reynolds and Mach number are kept constant at compressor inlet, an inevitable deviation in downstream flow causes mismatches in efficiency and pressure ratio. Both analytical and numerical methodologies show that the scale of shroud adjustments, as well as the size of mismatch in Mach and Reynolds number, can be correlated to the isentropic gas exponent. In summary, the main impact on gas behavior in an axial air compressor is attributable to the change in isentropic exponent. Derivations of shroud adaptation and analyses of inevitable aerodynamic mismatch are therefore developed depending on the isentropic exponent.

scholarly journals Ultra-Short Nacelles for Low Fan Pressure Ratio Propulsors

2014 ◽  
Author(s):  
Andreas Peters ◽  
Zoltán S. Spakovszky ◽  
Wesley K. Lord ◽  
Becky Rose
Keyword(s):  
Mach Number ◽  
Pressure Ratio ◽  
Design Tool ◽  
Operating Conditions ◽  
Design Parameters ◽  
Flow Distortion ◽  
Flow Area ◽  
Propulsive Efficiency ◽  
Flow Acceleration ◽  
Fuel Burn

As the propulsor fan pressure ratio (FPR) is decreased for improved fuel burn, reduced emissions and noise, the fan diameter grows and innovative nacelle concepts with short inlets are required to reduce their weight and drag. This paper addresses the uncharted inlet and nacelle design space for low-FPR propulsors where fan and nacelle are more closely coupled than in current turbofan engines. The paper presents an integrated fan-nacelle design framework, combining a spline-based inlet design tool with a fast and reliable body-force-based approach for the fan rotor and stator blade rows to capture the inlet-fan and fan-exhaust interactions and flow distortion at the fan face. The new capability enables parametric studies of characteristic inlet and nacelle design parameters with a short turn-around time. The interaction of the rotor with a region of high streamwise Mach number at the fan face is identified as the key mechanism limiting the design of short inlets. The local increase in Mach number is due to flow acceleration along the inlet internal surface coupled with a reduction in effective flow area. For a candidate short-inlet design with length over diameter ratio L/D = 0.19, the streamwise Mach number at the fan face near the shroud increases by up to 0.16 at cruise and by up to 0.36 at off-design conditions relative to a long-inlet propulsor with L/D = 0.5. As a consequence, the rotor locally operates close to choke resulting in fan efficiency penalties of up to 1.6 % at cruise and 3.9 % at off-design. For inlets with L/D < 0.25, the benefit from reduced nacelle drag is offset by the reduction in fan efficiency, resulting in propulsive efficiency penalties. Based on a parametric inlet study, the recommended inlet L/D is suggested to be between 0.25 and 0.4. The performance of a candidate short inlet with L/D = 0.25 was assessed using full-annulus unsteady RANS simulations at critical design and off-design operating conditions. The candidate design maintains the propulsive efficiency of the baseline case and fuel burn benefits are conjectured due to reductions in nacelle weight and drag compared to an aircraft powered by the baseline propulsor.

Sign in / Sign up

Export Citation Format

Share Document