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At least 343 records · Page 19

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NASA NTRSDate not supplied

Planar Investigation of a CobraMRV Reentry Flowfield Using Pulse-Burst, Cross-Correlation DGV

The flowfield around a Co-Optimization Blunt-body Reentry Analysis Mid-lift/drag-ratio Rigid Vehicle (CobraMRV) vehicle is investigated with a combination of high-speed planar laser Mie scattering (PLMS) and pulse-burst cross-correlation Doppler global velocimetry (PB-CC-DGV). Tests were conducted in the NASA Langley 4-foot Supersonic Unitary Plan Wind Tunnel (UPWT) over a range of different tunnel operating conditions and model configurations. Results indicate a complex shock-boundary layer interaction. Scalar information extracted from the PLMS show the evolution of the bow shock structure, while streamwise velocity measurements indicate the spatial evolution of the shock-boundary layer interaction including the growth of the separation shock foot and eventual reacceleration of the flow at farther downstream locations. Assessment of multiple cases show strong Mach and Reynolds numbers driven effects on the character of the shock-boundary layer interactions. Measurement uncertainties ranged from 50 to 150 m/s throughout the region of interest, driven largely by angular uncertainties and instabilities in the laser pointing. The mean accuracy of the freestream measurements was found to be 5.2-percent of tunnel predicted values.

Laser↗

NASA NTRS1988

Computations of vortical supersonic flows over a blunt-nose-cylinder at high angles of attack

Three-dimensional, viscous, and separated flows over a blunt-nose-cylinder at 20, 32, and 44-deg attack angles were computed. The approaching freestream was at a Mach number of 1.6 and a unit Reynolds number of 2 x 10 to the 6th/ft with a total temperature of 585 R. The cylinder used for the computations had a length-to-diameter ratio of 6.67 with a base diameter of 3 in. The flowfield was dominated by large-scale and multiple vortices generated by crossflow separation. The effect of turbulence on the flow structure of one case was modeled algebraically with modifications to correct the length and velocity scales in the regions of separation. The mass averaged Navier-Stokes equations were solved by an approximately factored, upwind-biased, implicit, finite volume scheme. The initialization of the flows was enhanced by a mesh sequencing strategy applied to the diagonalized form of the discretized equations. The convergence to steady-state was accelerated by a multigrid algorithm and using the block inversions for the discretized equations. Calculations were compared with experimental results.

Baysal, O.↗

NASA NTRS1976

Turbulence measurements in axisymmetric supersonic boundary layer flow in adverse pressure gradients

Measurements have been made of the mean-flow and turbulence properties in adiabatic turbulent boundary layer flows subjected to adverse pressure gradients. In the freestream region upstream of the adverse pressure gradient the Mach number was 3.86, the unit Reynolds number 5.3 million per foot. The boundary layer developed on the wall of an axisymmetric nozzle and straight test section. The pressure gradients at the test section wall were induced by contoured centerbodies mounted on the wind tunnel centerline. The flow under study simulated that which might be found in an axially symmetric engine inlet of a supersonic aircraft. The results obtained have shown good agreement to exist between the measured normalized turbulent velocity fluctuations and the results from other recent investigations of compressible boundary layers.

Gootzait, E.↗

NASA NTRS2002

An Evaluation of Parameters Influencing Jet Mixing Using the WIND Navier-stokes Code

The WIND code, a Reynolds-averaged Navier-Stokes solver used for a variety of aerospace flow simulations, was investigated for a Mach 2 nozzle at a series of nozzle stagnation temperatures. Comparisons of WIND calculations are made to experimental measurements of axial velocity, Mach number, and stagnation temperature along the jet centerline. The primary objective was to investigate the capabilities of the two-equation turbulence models available in WIND, version 4.0, for the analysis of heated supersonic nozzle flows. The models examined were the Menter Shear Stress Transport (SST) model and the Chien k-epsilon model, with and without the compressibility correction due to Sarkar. It was observed that all of the turbulence models investigated produced solutions that did not agree well with the experimental measurements. The effects of freestream Mach number and turbulent Prandtl number specifications were also investigated.

Dembowski, Mary Ann↗

NASA NTRS1991

Mach number validation of a new zonal CFD method (ZAP2D) for airfoil simulations

A closed-loop overlapped velocity coupling procedure has been utilized to combine a two-dimensional potential-flow panel code and a Navier-Stokes code. The fully coupled two-zone code (ZAP2D) has been used to compute the flow past a NACA 0012 airfoil at Mach numbers ranging from 0.3 to 0.84 near the two-dimensional airfoil C(lmax) point for a Reynolds number of 3 million. For these cases, the grid domain size can be reduced to 3 chord lengths with less than 3-percent loss in accuracy for freestream Mach numbers through 0.8. Earlier validation work with ZAP2D has demonstrated a reduction in the required Navier-Stokes computation time by a factor of 4 for subsonic Mach numbers. For this more challenging condition of high lift and Mach number, the saving in CPU time is reduced to a factor of 2.

Strash, Daniel J.↗

NASA NTRS1992

Numerical investigation of the effects of icing on fixed and rotary wing aircraft

A 2-D multi-element airfoil code was modified to study the effects of icing on the aerodynamic characteristics of high lift systems. In each zone of the flow field, the solver numerically integrates the 2-D compressible Navier-Stokes equations using a time marching scheme. The surface pressure distribution is generated over a GAW 130 airfoil/flap combination for a flap setting of 25 degrees, and an angle of attack equal to 5 degrees, at a freestream Mach number equal to 0.3. A series of calculations were performed to determine the effects of small scale ice build up on the high lift characteristics of this arifoil/flap combination. The appendix summarizes this progress. Joint studies on correlation of a 3-D iced wing code with experimental data reviewed new measured laser Doppler velocimeter data in the separated region behind the leading edge ice shape. A version of the iced wing analysis using the Roe scheme was developed to evaluate the poor correlation between the computed and measured velocities in the separated region. Work on the extension of the wing-alone analysis to wing body configuration began with modifications to the 3-D iced wing analyses to accept externally generated grids and multi-block grids.

Sankar, L. N.↗

NASA NTRSDate not supplied

Two-Line Focused Laser Differential Interferometry of a Flat Plate Boundary Layer at Mach 6

Flat plate models are often used in wind tunnel testing to make measurements of the boundary layer, but preclude the use of conventional focused laser differential interferometry (FLDI) measurements, particularly for measurements at the surface of the model. By instead focusing light in only one direction using cylindrical optics, as opposed to two directions using only spherical optics, FLDI measurements can be made without truncating the beam at the model edges. Measurements at discrete points along a line can also be acquired, effectively acting as many individual FLDI instruments in a single measurement. Here, a two-line FLDI system has been constructed that allows for the measurement lines to be easily oriented at any azimuthal angle about the optical axis since the optics are housed in rotation mounts. Velocity measurements at points along a line or between two lines are also possible, since the beams are split similar to the two-point FLDI instrument. The system was first characterized in the laboratory using a laser-induced breakdown spark to provide a well-defined density fluctuation, along with high-speed schlieren measurements for comparison. The system was then installed at the NASA Langley 20-Inch Mach 6 Air Tunnel to make measurements through the boundary layer of a flat plate model and in the tunnel freestream.

Joshua M Weisberger↗

NASA NTRS1996

Experimental Investigation of the Effects of Acceleration on Heat Transfer in the Turbulent Boundary Layer

The objective of this research was to experimentally investigate the combined effects of freestream acceleration and surface roughness on heat transfer and fluid flow in the turbulent boundary layer. The experiments included a variety of flow conditions ranging from aerodynamically smooth to transitionally rough to fully rough boundary layers with accelerations ranging from moderate to moderately strong. The test surfaces used were a smooth-wall test surface and two rough-wall surfaces which were roughened with 1.27 mm diameter hemispheres spaced 2 and 4 base diameters apart in a staggered array. The measurements consisted of Stanton number distributions, mean temperature profiles, skin friction distributions, mean velocity profiles, turbulence intensity profiles, and Reynolds stress profiles. The Stanton numbers for the rough-wall experiments increased with acceleration. For aerodynamically smooth and transitionally rough boundary layers, the effect of roughness is not seen immediately at the beginning of the accelerated region as it is for fully rough boundary layers; however, as the boundery layer thins under acceleration, the surface becomes relatively rougher resulting in a sharp increase in Stanton number.

Chakroun, Walid M.↗

NASA NTRS2005

Cold Flow Plume Entrainment Test Final Report NTF Test Number 2456

As part of the Space Shuttle Return to Flight (RTF) program, Marshall Space Flight Center (MSFC) performed computational fluid dynamics (CFD) analysis to define the velocity flowfields around the Shuttle stack at liftoff. These CFD predicted velocity flowfields were used in debris transport analysis (DTA). High speed flows such as plumes induce or 'entrain' mass from the surrounding environment. Previous work had shown that CFD analysis over-predicts plume induced flows. Therefore, the DTA would tend to 1) predict more debris impacts, and 2) the debris velocity (and kinetic energy) of those impacts would be too high. At a November, 2004 peer-review it was recommended that the Liftoff DTA team quantify the uncertainty in the DTA caused by the CFD's over prediction of plume induced flow. To do so, the Liftoff DTA team needed benchmark quality data for plume induced flow to quantify the CFD accuracy and its effect on the DTA. MSFC's Nozzle Test Facility (NTF) conducted the "Nozzle Induced Flows test, P#2456" to obtain experimental data for plume induced flows for nozzle flow exhausting into q quiescent freestream. Planning for the test began in December, 2004 and the experimental data was obtained in February and March of 2005. The funding for this test was provided by MSFC's Space Shuttle Propulsion Systems Integration and Engineering office.

Ruf, Joseph H.↗

DOE OSTI2023

Numerical Investigation of Wall-Cooling Effect on Aero-Optical Distortions for Hypersonic Boundary Layer

Compressible wall-modeled large-eddy simulations of Mach 8 turbulent boundary-layer flows over a flat plate were carried out for the conditions of the hypersonic wind tunnel at Sandia National Laboratories. The simulations provide new insight into the effect of wall cooling on the aero-optical path distortions for hypersonic turbulent boundary-layer flows. Four different wall-to-recovery temperature ratios, 0.3, 0.48, 0.71, and 0.89, are considered. Despite the much lower grid resolution, the mean velocity, temperature, and resolved Reynolds stress profiles from the simulation for a temperature ratio of 0.48 are in good agreement with those from a reference direct numerical simulation. The normalized root-mean-square optical path difference obtained from the present simulations is compared with that from reference direct numerical simulations, Sandia experiments, as well as predictions obtained with a semi-analytical model by Notre Dame University. Here the present analysis focuses on the effect of wall cooling on the wall-normal density correlations, on key underlying assumptions of the aforementioned model such as the strong Reynolds analogy, and on the elevation angle effect on the optical path difference. Wall cooling is found to increase the velocity fluctuations and decrease the density fluctuations, resulting in an overall reduction of the normalized optical path distortion. Compared to the simulations, the basic strong Reynolds analogy overpredicts the temperature fluctuations for cooled walls. Also different from the strong Reynolds analogy, the velocity and temperature fluctuations are not perfectly anticorrelated. Finally, as the wall temperature is raised, the density correlation length, away from the wall but inside the boundary layer, increases significantly for beam paths tilted in the downstream direction.

42 ENGINEERING↗

NASA NTRS1996

Emission Spectral Measurements in the Plenum of an Arc Jet Wind Tunnel

Arc jet wind tunnel facilities are used to evaluate thermal protection system materials for re-entry vehicles. The high speed, high temperature flowfield generated by the arc jet can simulate the extreme aerodynamic heating environment experienced during re-entry so that the survivability of heat shield materials and performance of various designs options can be tested. Although the re-entry heating environment can be approximated in the arc jet facility, the flowfield only partially simulates the actual re-entry flight conditions. Reynolds numbers are not matched so that surface shear stress distributions and mass transfer rates due to ablation or other mechanisms are not modeled correctly. Unlike flight conditions the arc freestream air is in non-equilibrium because of the rapid expansion that occurs in the supersonic nozzle. To properly study the actual re-entry flow environment, computational fluid dynamics, computational chemistry and radiation models must be used. Arc jet tunnel tests serve to validate these models. To perform accurate simulations inlet and boundary-conditions are needed, which come from measurements of the flowfield. The present study is concerned with measurements in the plenum region of an arc heater. In the past, conditions in the arc heater flowfield have been predicted using simulations since conventional measurement techniques could not be used in the harsh extremely high temperature environment. The present study is part of a recent push to utilize optical techniques to help better characterize the arc jet flowfields. Emission measurements have been made in the shock layer and the constrictor section of the arc heater to determine temperatures and species number densities. LIF measurements have been made in the free stream to determine temperature and velocity.

Donohue, Jim↗

NASA NTRSDate not supplied

Simulation and Modeling of Hypersonic Turbulent Boundary Layers Subject to Favorable Pressure Gradients due to Streamline Curvature

Direct numerical simulations (DNS) of favorable-pressure-gradient turbulent boundary layers are presented for a nominal freestream Mach number of 5, with the objective of assessing the limitations of the currently available Reynolds-averaged Navier-Stokes (RANS) models.The wall geometry and flow conditions of the DNS are representative of the experimental data for a Mach 4.9 turbulent boundary layer that was tested on a two-dimensional planar convexwall model in the high-speed blow-down wind tunnel located at the National Aerothermochemistry Laboratory at Texas A&M University. The favorable pressure gradient is induced bythe streamwise surface curvature. The DNS results will be compared with experimental measurements and RANS computations for the same flow conditions and wall geometry. The DNS datasets will also be analyzed to provide an improved physical understanding of hypersonic boundary-layer turbulence subject to mechanical nonequilibrium, including the validity of Morkovin’s hypothesis and the various available compressibility transformations, as well as to obtain turbulence statistics that are relevant to RANS modeling, including the distribution ofeddy viscosity, the budgets of turbulent kinetic energy and Reynolds stresses, and the velocity-temperature correlations. The RANS predictions will be obtained by using both the commonly used one- and two-equation models that assume a constant turbulent Prandtl number and an algebraic, variable Prandtl number model for the turbulent energy flux.

turbulent boundary layers↗

NASA NTRS2010

Boundary Layer Measurements in a Supersonic Wind Tunnel Using Doppler Global Velocimetry

A modified Doppler Global Velocimeter (DGV) was developed to measure the velocity within the boundary layer above a flat plate in a supersonic flow. Classic laser velocimetry (LV) approaches could not be used since the model surface was composed of a glass-ceramic insulator in support of heat-transfer measurements. Since surface flare limited the use of external LV techniques and windows placed in the model would change the heat transfer characteristics of the flat plate, a novel approach was developed. The input laser beam was divided into nine equal power beams and each transmitted through optical fibers to a small cavity within the model. The beams were then directed through 1.6-mm diameter orifices to form a series of orthogonal beams emitted from the model and aligned with the tunnel centerline to approximate a laser light sheet. Scattered light from 0.1-micron diameter water condensation ice crystals was collected by four 5-mm diameter lenses and transmitted by their respective optical fiber bundles to terminate at the image plane of a standard two-camera DGV receiver. Flow measurements were made over a range from 0.5-mm above the surface to the freestream at Mach 3.51 in steady state and heat pulse injected flows. This technique provides a unique option for measuring boundary layers in supersonic flows where seeding the flow is problematic or where the experimental apparatus does not provide the optical access required by other techniques.

Meyers, James F.↗

NASA NTRS2020

Simulation and Modeling of Cold-Wall Hypersonic Turbulent Boundary Layers on Flat Plate

Direct numerical simulations (DNS) of flat-plate, zero-pressure-gradient turbulent boundary layers are presented for nominal freestream Mach numbers of 11 and 14 and a highly cooled wall (wall-to-recovery temperature of approximately 0.2). The flow conditions of the DNS are representative of the experimental data for a Mach 11.1 turbulent boundary layer on a flat plate that was tested at Calspan–University of Buffalo Research Center (CUBRC) and the operational conditions of the AEDC Hypervelocity Tunnel No. 9 at Mach 14. The wall shear stress and turbulent heat flux predicted by DNS show good comparisons with those measured at CUBRC and those modeled with a Baldwin-Lomax model. The generated DNS database extends the range of Reynolds numbers from those typical of previous DNS studies for hypersonic boundary-layer flows and confirms the validity of Morkovin’s hypothesis within the relatively broader range of Reynolds numbers considered in this study. A priori assessment of turbulent heat flux models with DNS data shows that the algebraic energy flux (AEF) formulation gives better predictions of the streamwise turbulent heat flux than the commonly used constant turbulent Prandtl number model while both models give similarly good predictions of the wall-normal turbulent heat flux. A posteriori assessment shows that the AEF model improves the prediction of the temperature profiles while keeping a good prediction of the velocity field.

Junji Huang↗

NASA NTRS2002

Synthetic Jet Interaction With A Turbulent Boundary Layer Flow

Perhaps one of the more notable advances to have occurred in flow control technology in the last fifteen years is the application of surface-issuing jets for separation control on aerodynamic surfaces. The concept was introduced by Johnston and Night (1990) who proposed using circular jets, skewed and inclined to the wall, to generate streamwise vortices for the purpose of mitigating boundary layer separation. The skew and inclination angles have subsequently been shown to affect the strength and sign of the ensuing vortices. With a non-circular orifice, in addition to skew and inclination, the yaw angle of the major axis of the orifice can influence the flow control effectiveness of the jet. In particular, a study by Chang arid Collins (1997) revealed that a non-circular orifice, yawed relative to the freestream, can be used to control the size and strength of the vortices produced by the control jet. This early work used jets with only a steady injection of mass. Seifert et al. revealed that an unsteady blowing jet, could be as effective at separation control as a steady jet but with less mass flow. Seifert et al. showed that small amplitude blowing oscillations superimposed on a low momentum steady jet Was the most effective approach to delaying separation on a NACA 0015 airfoil at post-stall angles of attack. More recent work suggests that perhaps the most efficient jet control effect comes from a synthetic (oscillatory) jet where the time-averaged mass flux through the orifice is zero, but the net wall normal momentum is non-zero. The control effectiveness of synthetic jets has been demonstrated for several internal and external flow fields used synthetic jet control on a thick, blunt-nosed airfoil to delay stall well beyond the stall angles for the uncontrolled airfoil and with a dramatic increase in the lift-to-drag performance. Amitay et al. used an array of synthetic jets to mitigate flow separation in curved and diffusing ducts. While the control effectiveness of synthetic jet actuators in an application setting has been confirmed through a number of studies, the fluid dynamic mechanism by which control is effected is not well understood. Issues like yaw angle arid velocity ratio that have been studied for steady control jets have not been investigated for synthetic jets. Moreover, the role played by the inherently unsteady nature of a synthetic jet in the interaction with the controlled flow is not known. Some recent work by Rinehart and Glazer and Smith suggests that away from tile immediate vicinity of the jet orifice the flow field is steady in the mean, and that the interaction of a synthetic jet and a boundary layer creates a secondary flow in the boundary layer consisting primarily of streamwise vortices.

Smith, Douglas R.↗

NASA NTRS2022

Nonlinear Evolution of Instabilities in a Laminar Separation Bubble at a Hypersonic Mach Number

The development of both convective stationary perturbation as well as global instabilities in the vicinity of a laminar separation bubble above an axisymmetric compression corner in a hypersonic flow is investigated using numerical simulations. The flow configuration of primary interest corresponds to the cone-cylinder-flare model used in experimental measurements in the Boeing/AFOSR Mach-6 Quiet Tunnel at Purdue University. For a flare angle of 10 degrees and a unit Reynolds number of 11.5 x 10 6 m -1 , their surface flow visualizations identified the presence of streamwise elongated thermal streaks near the reattachment position and the dominant azimuthal spacing between the streaks was determined to be approximately 10 degrees (i.e., an azimuthal mode number of 36). Previous linear stability analyses predicted that the amplification characteristics of small amplitude, unsteady, convective instabilities within this flow were consistent with the surface pressure fluctuations measured in the experiment. However, their accompanying analysis of global instabilities had found the separation bubble to be weakly unstable at the 10 degrees flare angle, with the most unstable global mode corresponding to a stationary disturbance with an azimuthal wavenumber of m ≈ 5, which was well below the measured wavenumber of m = 36. Besides confirming the presence of the global instability at these flow conditions, the present numerical simulations quantify the details of the stationary equilibrium state associated with the supercritical bifurcation resulting from the nonlinear saturation of the unstable global mode. Although velocity perturbations associated with the saturated global mode are dominated by the fundamental spanwise wavelength associated with the linear global instability, the surface heat flux downstream of reattachment is dominated by m = 36 in agreement with experimental measurements. However, the shorter wavelength peak ( m = 36) in the azimuthal spectrum is not manifested as streaks with an azimuthal spacing of 10 degrees. An additional simulation of the combined evolution of both unsteady convective and global instabilities indicates that the nonlinear interactions between a broadband spectrum of oblique Mack’s first mode instabilities generate stationary vortex modes with a higher dominant wavenumber ( m = 60) as that observed in the experiments. Additional analyses are performed to predict the global mode behavior at other flare angles and how it differs from the convective centrifugal instability due to concave streamline curvature over the compression corner.

Hypersonic flow↗

NASA NTRSDate not supplied

Wall-modeled LES of the NASA Juncture Flow Experiment

We perform wall-modeled large eddy simulation (WMLES) of the NASA Juncture Flow experiment (Kegerise & Neuhart, NASA/TM–2019–20286) using an equilibrium wall model and unstructured finite volume solver, to assess its predictive ability for complex geometries. The flow condition simulated corresponds to 5 degrees angle of attack with a freestream Mach number of 0.189, and a Reynolds number based on the mean crank chord of 2.4 million. All simulations neglect the effects of wind tunnel walls, sting and mast present in the experiment. To assess the performance of WMLES on simpler flows with very coarse grids, we simulate flow in a turbulent channel at a friction Reynolds number, Reτ≈2000, and flow through a square duct at anReτ≈1000. The duct flow results using the equilibrium wall model indicate that the skin friction is inaccurate near the corner for grid topologies and resolutions typically used in WMLES, having implications for more complicated flows such as the juncture flow. For the juncture flow simulation, two different approaches are investigated. The first approach consists of a truncated-domain simulation wherein the inflow plane of the computational domain is placed at≈0.2chord length of the wing, and we prescribe the mean flow from a separate Reynolds-averaged Navier-Stokes (RANS) solution along with synthetic turbulence to initiate realistic unsteadiness in the domain. The second approach involves simulating the entire geometry with trips to trigger transition to turbulence. The truncated- and full-domain simulations contain about 62 and 90 million cells, respectively; with 8-10 points per boundary layer thickness in the wall-parallel directions, 14-20 points in the wall-normal direction and a near-wall viscous spacing (∆n+1) of≈100. Preliminary results are encouraging overall in terms of the prediction of wall pressure, wall skin friction, velocity and stresses; but indicate that further work is required to improve the predictions in the separation bubble and wing-fuselage corner regions. The grid resolution used in this study is still fairly coarse, and the results should be interpreted as work in progress.

Prahladh S Iyer↗

NASA NTRS2009

Turboelectric Distributed Propulsion Engine Cycle Analysis for Hybrid-Wing-Body Aircraft

Meeting NASA's N+3 goals requires a fundamental shift in approach to aircraft and engine design. Material and design improvements allow higher pressure and higher temperature core engines which improve the thermal efficiency. Propulsive efficiency, the other half of the overall efficiency equation, however, is largely determined by the fan pressure ratio (FPR). Lower FPR increases propulsive efficiency, but also dramatically reduces fan shaft speed through the combination of larger diameter fans and reduced fan tip speed limits. The result is that below an FPR of 1.5 the maximum fan shaft speed makes direct drive turbines problematic. However, it is the low pressure ratio fans that allow the improvement in propulsive efficiency which, along with improvements in thermal efficiency in the core, contributes strongly to meeting the N+3 goals for fuel burn reduction. The lower fan exhaust velocities resulting from lower FPRs are also key to meeting the aircraft noise goals. Adding a gear box to the standard turbofan engine allows acceptable turbine speeds to be maintained. However, development of a 50,000+ hp gearbox required by fans in a large twin engine transport aircraft presents an extreme technical challenge, therefore another approach is needed. This paper presents a propulsion system which transmits power from the turbine to the fan electrically rather than mechanically. Recent and anticipated advances in high temperature superconducting generators, motors, and power lines offer the possibility that such devices can be used to transmit turbine power in aircraft without an excessive weight penalty. Moving to such a power transmission system does more than provide better matching between fan and turbine shaft speeds. The relative ease with which electrical power can be distributed throughout the aircraft opens up numerous other possibilities for new aircraft and propulsion configurations and modes of operation. This paper discusses a number of these new possibilities. The Boeing N2 hybrid-wing-body (HWB) is used as a baseline aircraft for this study. The two pylon mounted conventional turbofans are replaced by two wing-tip mounted turboshaft engines, each driving a superconducting generator. Both generators feed a common electrical bus which distributes power to an array of superconducting motor-driven fans in a continuous nacelle centered along the trailing edge of the upper surface of the wing-body. A key finding was that traditional inlet performance methodology has to be modified when most of the air entering the inlet is boundary layer air. A very thorough and detailed propulsion/airframe integration (PAI) analysis is required at the very beginning of the design process since embedded engine inlet performance must be based on conditions at the inlet lip rather than freestream conditions. Examination of a range of fan pressure ratios yielded a minimum Thrust-specific-fuel-consumption (TSFC) at the aerodynamic design point of the vehicle (31,000 ft /Mach 0.8) between 1.3 and 1.35 FPR. We deduced that this was due to the higher pressure losses prior to the fan inlet as well as higher losses in the 2-D inlets and nozzles. This FPR is likely to be higher than the FPR that yields a minimum TSFC in a pylon mounted engine. 1

Felder, James L.↗