A trade-factor-based system study has been carried out to identify fuel burn benefits associated with boundary layer ingestion (BLI) for generation-after-next (N+2) aircraft and propulsion system concepts. The analysis includes detailed propulsion system engine cycle modeling for a next-generation, Ultra-High-Bypass (UHB) propulsion system with BLI using the Numerical Propulsion System Simulation (NPSS) computational model. Cycle modeling was supplemented with one-dimensional theory to identify limiting theoretical BLI benefits associated with the blended wing body reference vehicle used in the study. The system study employed low-order models of engine extractions associated with inlet flow control; nacelle weight and drag; fan performance; and inlet pressure losses. Aircraft trade factors were used to estimate block fuel burn reduction for a long-range commercial transport mission. Results of the study showed that a 3-5% BLI fuel burn benefit can be achieved for N+2 aircraft relative to a baseline high-performance, pylon-mounted, UHB propulsion system. High-performance, distortion-tolerant turbomachinery, and low-loss, low-drag inlet systems, were identified as key enabling technologies. Larger benefits were estimated for N+3 configurations for which larger fractions of aircraft boundary layer can be ingested. NomenclatureA = area (in. 2 ) AR = inlet aspect ratio (w / h) c, C = aircraft chord (ft or in.) D = amount of aircraft viscous drag ingested by propulsion systems (lbf) F n , F N = engine net thrust (lbf) FB = fuel burn (lbs) h = inlet height (ft or in.) H = boundary layer shape factor (δ* / θ) k = boundary layer pseudo-energy thickness (in.) K = boundary layer pseudo-energy factor (k / θ) M = Mach number n = unit surface vector 2 P = pressure (psi) P T , P t = total pressure (psi) R = wake recovery factor (1-Δ j /Δ 0 ) T = thrust (lbf); temperature (°R) U = velocity (ft / s) V = free stream velocity (ft / s) V x = axial component of free stream velocity (ft / s) w, W = inlet width (ft or in.) x, X = axial coordinate or dimension (in.) y = transverse or vertical coordinate or dimension (in.)Greek: δ* = boundary layer displacement thickness (in.) Δ = wake velocity defect relative to freestream or jet velocity condition ρ = density (slug / ft 3 ) τ = wall shear stress (psf) θ = boundary layer momentum thickness (in.) Subscripts / Superscripts:∞, 0 = freestream condition j = propulsion system jet velocity condition MA = mass averaged quantity s = static condition T = stagnation condition x = axial component

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The paper describes the aerodynamic CFD analysis that was conducted to address the integration of an embedded-engine (EE) inlet with the fan stage. A highly airframe-integrated, distortion-tolerant propulsion preliminary design study was carried out to quantify fuel burn benefits associated with boundary layer ingestion (BLI) for “N+2” blended wing body (BWB) concepts. The study indicated that low-loss inlets and high-performance, distortion-tolerant turbomachines are key technologies required to achieve a 3–5% BLI fuel burn benefit relative to a baseline high-performance, pylon-mounted, propulsion system. A hierarchical, multi-objective, computational fluid dynamics-based aerodynamic design optimization that combined global and local shaping was carried out to design a high-performance embedded-engine inlet and an associated fan stage. The scaled-down design will be manufactured and tested in NASA’s 8′×6′ Transonic Wind Tunnel. Unsteady calculations were performed for the coupled inlet and fan rotor and inlet, fan rotor and exit guide vanes. The calculations show that the BLI distortion propagates through the fan largely un-attenuated. The impact of distortion on the unsteady blade loading, fan rotor and fan stage efficiency and pressure ratio is analyzed. The fan stage pressure ratio is provided as a time-averaged and full-wheel circumferential-averaged value. Computational analyses were performed to validate the system study and design-phase predictions in terms of fan stage performance and operability. For example, fan stage efficiency losses are less than 0.5–1.5% when compared to a fan stage in clean flow. In addition, these calculations will be used to provide pretest predictions and guidance for risk mitigation for the wind tunnel test.

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The test section of the 8-by 6-Foot Supersonic Wind Tunnel at NASA Glenn Research Center was modified to produce the test conditions for a boundary-layer-ingesting propulsor. A test was conducted to measure the flow properties in the modified test section before the propulsor was installed. Measured boundary layer and freestream conditions were compared to results from computational fluid dynamics simulations of the external surface for the reference vehicle. Testing showed that the desired freestream conditions and boundary layer thickness could be achieved; however, some non-uniformity of the freestream conditions, particularly the total temperature, were observed. Nomenclature 86 SWT 8-by 6-Foot Supersonic Wind Tunnel BL Boundary Layer BLI Boundary Layer Ingesting BWB Blended Wing Body CFD Computational Fluid Dynamics TS Tunnel Station, in. h height of the boundary layer thickening pins, in. u velocity in the streamwise direction, ft/s U ∞ freestream velocity in the streamwise direction, ft/s x axial coordinate, tunnel station, in. y transverse coordinate, in. z vertical coordinate, in.  boundary layer thickness (99 percent of the freestream velocity), in.