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The tests described in this paper afford a direct comparison of the efficiency and smoothness of flow obtained with propeller fan and multiblade centrifugal fan drives in the same wind tunnel. The propeller fan was found to be superior to the centrifugal fan in that the efficiency was about twice as great, and the flow much smoother.
Limited data indicates the feasibility of designing fan stages for operation with weak oblique shocks in the rotor blade tip region. A 1600 ft/sec rotor tip speed fan indicated an overall efficiency of 0.846 at a pressure ratio of 1.51 even though the rotor blades had incurred damage in the leading edge tip region prior to obtaining the design speed performance. After the test it was observed that a number of the rotor blade mid-span dampers had failed, subsequently damaging a number of rotor blades in the leading edge tip region.
The NASA two-stage fan was tested with a redesigned first-stage rotor. The redesign included a new design approach to account for the presence of a part-span damper. At design speed the fan achieved a peak efficiency of 0.803, which is 1.9 percentage points higher than the original design. The peak efficiencies of the first stage and first rotor were 0.789 and 0.821, respectively. An improvement in efficiency of up to 5 percentage points in the damper region was achieved over the original large damper version. The stall margin, based on flow conditions at peak efficiency, was 10 percent at design speed.
The Dynamic Spin Rig Facility at the NASA Glenn Research Center is used to determine the structural response of rotating engine components without the effects of aerodynamic loading. Recently, this rig's capabilities were enhanced through the replacement of grease-lubricated ball bearings with magnetic bearings. Magnetic bearings offer a number of advantages--the most important here being that they not only fully support the rotor system, but excite it as well. Three magnetic bearings support the rotor and provide five axes of controlled motion: an x- and y-axis translation at each of two radial bearings and a z-axis translation in the vertical or axial direction. Sinusoidal excitation (most commonly used) can be imparted on the rotor through the radial magnetic bearings in either a fixed or rotating frame of reference. This excitation is added directly to the magnetic bearing control output. Since the rotor is fully levitated, large translations and rotations of the rotor system can be achieved. Some of the capabilities of this excitation system were determined and reported. The accelerations obtained at the tip of a titanium flat plate test article versus the swept sine excitation sent to both radial bearings in phase and perpendicular to the plane containing the two blades are shown. Recent tests required the excitation of fundamental bending and torsional blade resonances at rotor speeds up to 10,000 rpm. Successful fixed synchronous rotation of the excitation signal provided the best detectable blade resonant vibrations at excitation frequencies up to 1100 Hz for the particular blades of interest. A noncontacting laser measurement system was used to collect blade-tip motions. From these data, the amplitude and frequency of the motion could be determined as well as the blade damping properties. Damping could be determined using two methods: (1) free decay and (2) curve fitting the vibration amplitude as a function of frequency in and around the resonance of interest and using the half-power method. The free decay of a composite blade vibrating at its first bending resonance while rotating at 3000 rp is shown. This new system is currently being used to support the Efficient Low-Noise Fan project at Glenn. The damping properties of prototype hollow composite blades specially designed to reduce fan noise are currently being determined.
The SUbsonic Single Aft eNgine (SUSAN) program is being developed as a sustainable subsonic regional aircraft. NASA is conducting trade studies for this future hybrid single aisle passenger aircraft. The goal of reducing emission levels by 50% will be enabled through an advanced design. An aft mounted single turbine engine provides thrust and electric power to drive 16 wing mounted electric engines. Using the 16 counter-rotating electric fans will create an efficient, very high bypass ratio, engine. Electric engines allow for distributed thrust, reducing control surfaces and provide a greater dynamic response. Batteries provide boost for takeoff and climb, allowing a smaller, more efficient, turbine engine. As part of the SUSAN effort, two pilot-in-the-loop studies were conducted at NASA Langley Research Center in Virginia to explore flight deck design issues. In particular: how many throttles are required/desired for a 17-engine aircraft, does the hybrid electric design change the answer, what are the requirements for electric engine displays, what system pages are required, and what information is required to monitor the batteries. Objective and subjective data were collected after each scenario and at the end of the day with a guided set of questionnaires from sixteen crews paired by airline. Data along with pilot comments were used to improve the throttle and engine displays between the studies and used to make the final recommendations on flight deck configurations. This paper provides final design recommendations for the SUSAN flight deck and justification for each recommendation.
A single-stage axial fan was built and tested in the shop of the propeller-research tunnel of the NACA. The fan comprised a simple 24-blade rotor having a diameter of 21 inches and a solidity of 0.86 and a set of 37 contravanes having a solidity of 1.33. The rotor was driven by a 25-horsepower motor capable of rotating at a speed of 3600 r.p.m. The fan was tested for volume, pressure, and efficiency over a range of delivery pressures and volumes for a wide range of contravane and blade-angle settings. The test results are presented in chart form in terms of nondimensional units in order that similar fans may be accurately designed with a minimum effort. The maximum efficiency (88 percent) was obtained by the fan at a blade angle of 30 degrees and a contravane angle of 70 degrees. An efficiency of 80 percent was obtained by the fan with the contravanes removed.
A graphical presentation of the aerodynamic data acquired during coannular nozzle performance wind tunnel tests is given. The graphical data consist of plots of nozzle gross thrust coefficient, fan nozzle discharge coefficient, and primary nozzle discharge coefficient. Normalized model component static pressure distributions are presented as a function of primary total pressure, fan total pressure, and ambient static pressure for selected operating conditions. In addition, the supersonic cruise configuration data include plots of nozzle efficiency and secondary-to-fan total pressure pumping characteristics. Supersonic and subsonic cruise data are given.
The Glenn Flux Reconstruction (GFR) code is a computational fluid dynamics (CFD) code under development at NASA Glenn Research Center. GFR is based on the high-order flux reconstruction (FR) method and provides a large-eddy simulation (LES) capability that is both accurate and efficient for complex aeropropulsion flows. Three significant new capabilities have been added to the code that improve its performance and functionality. First, a variety of explicit Runge-Kutta methods, including some with adaptive time stepping, were added to GFR with two methods offering a 33% improvement in time-to-solution. Second, GFR can now utilize fully unstructured, mixed-element meshes to more easily facilitate the grid generation process for complex geometries. Finally, a rotating reference frame capability has been added to GFR for solving rotating turbomachinery problems. A selection of results demonstrating these new capabilities are presented in this work. The Taylor-Green vortex problem is used to verify the new unstructured capability by showing similar accuracy and resolution for all element types. LES of the Turbulent Heat Flux Phase III (THX3) experiment with comparison to another high-order LES code and a popular Reynolds-averaged Navier-Stokes (RANS) code demonstrates the accuracy of the code for complex aeropropulsion flows. Finally, LES of a spacecraft cabin ventilation fan shows the ability of GFR to efficiently establish a fan performance map and identify operating points for further analysis at high orders of accuracy.
The Glenn Flux Reconstruction (GFR) code is a computational fluid dynamics (CFD) code under development at NASA Glenn Research Center. GFR is based on the high-order flux reconstruction (FR) method and provides a large-eddy simulation (LES) capability that is both accurate and efficient for complex aeropropulsion flows. Three significant new capabilities have been added to the code that improve its performance and functionality. First, a variety of explicit Runge-Kutta methods, including some with adaptive time stepping, were added to GFR with two methods offering a 33% improvement in time-to-solution. Second, GFR can now utilize fully unstructured, mixed-element meshes to more easily facilitate the grid generation process for complex geometries. Finally, a rotating reference frame capability has been added to GFR for solving rotating turbomachinery problems. A selection of results demonstrating these new capabilities are presented in this work. The Taylor-Green vortex problem is used to verify the new unstructured capability by showing similar accuracy and resolution for all element types. LES of the Turbulent Heat Flux Phase III (THX3) experiment with comparison to another high-order LES code and a popular Reynolds-averaged Navier-Stokes (RANS) code demonstrates the accuracy of the code for complex aeropropulsion flows. Finally, LES of a spacecraft cabin ventilation fan shows the ability of GFR to efficiently establish a fan performance map and identify operating points for further analysis at high orders of accuracy.
With propulsion research programs focused on new levels of efficiency and noise reduction, two avenues for advanced gas turbine technology are emerging: the geared turbofan and ultrahigh bypass ratio fan engines. Both of these candidates are being pursued as collaborative research projects between NASA and the engine manufacturers. The high bypass concept from GE Aviation is an unducted fan that features a bypass ratio of over 30 along with the accompanying benefits in fuel efficiency. This project improved the test and measurement capabilities of the unducted fan blade dynamic response. In the course of this project, Mechanical Solutions, Inc. (MSI) collaborated with GE Aviation to (1) define the requirements for fan blade measurements; (2) leverage MSI's radar-based system for compressor and turbine blade monitoring; and (3) develop, validate, and deliver a noncontacting blade vibration measurement system for unducted fans.
A Turboelectric Distributed Propulsion (TeDP) system differs from other propulsion systems by the use of electrical power to transmit power from the turbine to the fan. Electrical power can be efficiently transmitted over longer distances and with complex topologies. Also the use of power inverters allows the generator and motors speeds to be independent of one another. This decoupling allows the aircraft designer to place the core engines and the fans in locations most advantageous for each. The result can be very different installation environments for the different devices. Thus the installation effects on this system can be quite different than conventional turbofans where the fan and core both see the same installed environments. This paper examines a propulsion system consisting of two superconducting generators, each driven by a turboshaft engine located so that their inlets ingest freestream air, superconducting electrical transmission lines, and an array of superconducting motor driven fan positioned across the upper/rear fuselage area of a hybrid wing body aircraft in a continuous nacelle that ingests all of the upper fuselage boundary layer. The effect of ingesting the boundary layer on the design of the system with a range of design pressure ratios is examined. Also the impact of ingesting the boundary layer on off-design performance is examined. The results show that when examining different design fan pressure ratios it is important to recalculate of the boundary layer mass-average Pt and MN up the height for each inlet height during convergence of the design point for each fan design pressure ratio examined. Correct estimation of off-design performance is dependent on the height of the column of air measured from the aircraft surface immediately prior to any external diffusion that will flow through the fan propulsors. The mass-averaged Pt and MN calculated for this column of air determine the Pt and MN seen by the propulsor inlet. Since the height of this column will change as the amount of air passing through the fans change as the propulsion system is throttled, and since the mass-average Pt and MN varies by height, this capture height must be recalculated as the airflow through the propulsor is varied as the off-design performance point is converged.
A computationally efficient Boundary Integral Equation Method (BIEM) for the prediction of ducted fan engine noise is presented. The key features of the BIEM are its versatility and the ability to compute rapidly any portion of the sound field without the need to compute the entire field. Governing equations for the BIEM are based on the assumptions that all acoustic processes are linear, generate spinning modes, and occur in a uniform flow field. An exterior boundary value problem (BVP) is defined that describes the scattering of incident sound by an engine duct with arbitrary profile. Boundary conditions on the duct walls are derived that allow for passive noise control treatment. The BVP is recast as a system of hypersingular boundary integral equations for the unknown duct surface quantities. BIEM solution methodology is demonstrated for the scattering of incident sound by a thin cylindrical duct with hard walls. Numerical studies are conducted for various engine parameters and continuous portions of the total pressure field are computed. Radiation and duct propagation results obtained are in agreement with the classical results of spinning mode theory for infinite ducts.
Increased need for more efficient long-range supersonic flight has revived interest in the supersonic throughflow fan as a possible component for advanced high-speed propulsion systems. A fan that can operate with supersonic inlet axial Mach numbers would reduce the inlet losses incurred in diffusing the flow from supersonic Mach numbers to a subsonic one at the fan face. In addition, the size and weight of an all-supersonic inlet will be substantially lower than those of a conventional inlet. However, the data base for components of this type is practically nonexistent. Therefore, in order to furnish the required information for assessing the potential for this type of fan, the NASA Lewis Research Center has begun a program to design, analyze, build, and test a fan stage that is capable of operating with supersonic axial velocities from inlet to exit. The objectives are to demonstrate the feasibility and potential of supersonic throughflow fans, to gain a fundamental understanding of the flow physics associated with such systems, and to develop an experimental data base for design and analysis code validation. A brief overview of past supersonic throughflow fan activities are provided; the technology needs discussed; the design of a supersonic throughflow fan stage, a facility inlet, and a downstream diffuser described; and the results from the analysis codes used in executing the design are presented. Also presented is an engine concept intended to permit establishing supersonic throughflow within the fan on the runway and maintaining the supersonic throughflow condition within the fan throughout the flight envelope.
A design technology study was performed to identify a high speed, multistage, variable geometry fan configuration capable of achieving wide flow modulation with near optimum efficiency at the important operating condition. A parametric screening study of the front and rear block fans was conducted in which the influence of major fan design features on weight and efficiency was determined. Key design parameters were varied systematically to determine the fan configuration most suited for a double bypass, variable cycle engine. Two and three stage fans were considered for the front block. A single stage, core driven fan was studied for the rear block. Variable geometry concepts were evaluated to provide near optimum off design performance. A detailed aerodynamic design and a preliminary mechanical design were carried out for the selected fan configuration. Performance predictions were made for the front and rear block fans.
In order to meet the emissions and efficiency goals established by the aviation community, next generation aircraft engines will require low pressure ratio, distortion tolerant fans. Maturation of the necessary technologies will require facilities which are capable of accurately testing the fan performance. To achieve this, the Single Stage Axial Compressor and Fan facility at NASA Glenn Research Center was upgraded. Details of the goals and upgrades undertaken were documented in the 2022 Turbo Expo paper, “Upgrades to the Single Stage Axial Compressor and Fan Facility for Low Pressure Ratio Boundary Layer Ingesting Fan Research.” The performance of the upgraded facility will be documented in this paper. This paper will review the challenges associated with testing low pressure ratio and distortion tolerant fans and briefly review the upgrades undertaken and their goals. Then the performance of the facility upgrades and the extent to which they were able to achieve the stated goals will be discussed. This includes mass flow measurements, test section temperature, pressure, flow angle at two axial locations, and turbulence levels at the inlet and test section. The paper also documents the performance of distortion generating swirl vanes and total pressure screens to achieve desired distorted conditions at the Aerodynamic Interface Plane.
In order to meet the emissions and efficiency goals established by the aviation community, next generation aircraft engines will require low pressure ratio, distortion tolerant fans. Maturation of the necessary technologies will require facilities capable of accurately testing the fan performance are needed. To achieve this, the Single Stage Axial Compressor and Fan facility at NASA Glenn Research Center was upgraded. Details of the goals and upgrades undertaken were documented in the 2022 Turbo Expo paper, “Upgrades to the Single Stage Axial Compressor and Fan Facility for Low Pressure Ratio Boundary Layer Ingesting Fan Research.” The performance of the upgraded facility will be documented in this paper. This paper will review the challenges associated with testing low pressure ratio and distortion tolerant fans and briefly review the upgrades undertaken and their goals. Then the performance of the facility upgrades and the extent to which they were able to achieve the stated goals. This includes mass flow measurements, test section temperature, pressure, flow angle at two axial locations, and turbulence levels at the inlet and test section. The paper also documents the performance of distortion generating swirl vanes and total pressure screens to achieve desired distorted conditions at the Aerodynamic Interface Plane.
A fast-running unsteady aerodynamics code, LINFLUX, was previously developed for predicting turbomachinery flutter. This linearized code, based on a frequency domain method, models the effects of steady blade loading through a nonlinear steady flow field. The LINFLUX code, which is 6 to 7 times faster than the corresponding nonlinear time domain code, is suitable for use in the initial design phase. Earlier, this code was verified through application to a research fan, and it was shown that the predictions of work per cycle and flutter compared well with those from a nonlinear time-marching aeroelastic code, TURBO-AE. Now, the LINFLUX code has been applied to real configurations: fans developed under the Energy Efficient Engine (E-cubed) Program and the Quiet Aircraft Technology (QAT) project. The LINFLUX code starts with a steady nonlinear aerodynamic flow field and solves the unsteady linearized Euler equations to calculate the unsteady aerodynamic forces on the turbomachinery blades. First, a steady aerodynamic solution is computed for given operating conditions using the nonlinear unsteady aerodynamic code TURBO-AE. A blade vibration analysis is done to determine the frequencies and mode shapes of the vibrating blades, and an interface code is used to convert the steady aerodynamic solution to a form required by LINFLUX. A preprocessor is used to interpolate the mode shapes from the structural dynamics mesh onto the computational fluid dynamics mesh. Then, LINFLUX is used to calculate the unsteady aerodynamic pressure distribution for a given vibration mode, frequency, and interblade phase angle. Finally, a post-processor uses the unsteady pressures to calculate the generalized aerodynamic forces, eigenvalues, an esponse amplitudes. The eigenvalues determine the flutter frequency and damping. Results of flutter calculations from the LINFLUX code are presented for (1) the E-cubed fan developed under the E-cubed program and (2) the Quiet High Speed Fan (QHSF) developed under the Quiet Aircraft Technology project. The results are compared with those obtained from the TURBO-AE code. A graph of the work done per vibration cycle for the first vibration mode of the E-cubed fan is shown. It can be seen that the LINFLUX results show a very good comparison with TURBO-AE results over the entire range of interblade phase angle. The work done per vibration cycle for the first vibration mode of the QHSF fan is shown. Once again, the LINFLUX results compare very well with the results from the TURBOAE code.
A limited study has been conducted to establish the performance and noise characteristics of a low design tip speed (168 m/s, 550 ft/sec) low pressure ratio (1.04) variable pitch fan which was tested in the Langley 30 X 60 tunnel. This fan was designed for minimum noise when installed in the tail mount location of a twin engine aircraft which normally has both nose and tail mounted propulsors. Measurements showed the fan noise to be very close to predictions made during the design of the fan and extremely low in level (65 dBA at 1000 ft) with no acoustic treatment. This is about 8 dB lower than the unshrouded 2 blade propeller normally used in this installation. On the basis of tests conducted during this program, it appears that this level could be further reduced by 2 dBA if optimized acoustic treatments were installed in the fan duct. Even the best of the shrouded propellers tested previously were 7 dB higher in level than the Q-Fan without acoustic treatment. It was found that the cruise performance of this fan was within 5% of the predicted efficiency of 72%. Evaluation of the performance data indicated that disturbances in the inflow to the fan were the probable cause of the reduced performance.