AIAA 2014-0907 AIAA SciTech 13-17 January 2014, National Harbor, Maryland 52nd Aerospace Sciences Meeting
Computational Assessment of the Boundary Layer Ingesting Nacelle Design of the D8 Aircraft Shishir A. Pandya∗ NASA Ames Research Center, Moffett Field, CA, USA
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Arthur Huang†, Alejandro Espitia‡, and Alejandra Uranga§ Massachusettes Institute of Technology, Cambridge, MA, USA The Overflow viscous overset solver with an actuator-disk propulsor model is used to simulate the flow around a next-generation subsonic transport concept, the D8 aircraft, and demonstrate the benefit of BLI for this configuration. Results from an unpowered version of the aircraft are validated against a wind tunnel experiment, followed by a comparison of the performance of two D8 configurations: one with conventional, non-BLI, propulsors, and one with BLI propulsors. The BLI configuration is shown to require 9% less propulsor power relative to the conventional configuration at the simulated cruise condition. It is shown that the difference is due to a decrease in entropy generation, associated with wake and jet mixing. Keywords: BLI, boundary-layer ingestion, Overset grids, Aerodynamics, D8, Doublebubble aircraft.
I.
Introduction
The goal of the Fixed Wing (FW) Project in NASA’s Fundamental Aeronautics Program is to develop technologies and concepts to drastically improve the energy efficiency and reduce the environmental impact of future commercial transport aircraft. The project focuses on the ‘N+3’ timeframe (i.e. three generations beyond the current) with notional entry into service in approximately 2030-35. The project is leveraging both in-house and external expertise. The goal Figure 1. The MIT D8 “double-bubble” aicraft concept. of the FW ‘N+3’ solicitation is to stimulate thinking to address pressing concerns related to commercial aviation, such as reducing energy consumption, environmental impact, noise, as well as dealing with future operations challenges. To achieve this, the project is identifying advanced airframe and propulsion systems concepts, and is working to bring enabling technologies to maturity. ∗ Aerospace
Engineer, NASA Advasnced Supercomputing Division, Mail Stop 258-2, AIAA Senior Member. Engineer, Dept. of Aeronautics and Astronautics. ‡ Former M.S. student, Dept. of Aeronautics and Astronautics, AIAA Student Member. § Research Engineer, Dept. of Aeronautics and Astronautics,
[email protected], AIAA Member. † Research
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This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.
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This paper presents a study of the effect of a boundary layer ingesting (BLI) propulsor on the aerodynamic performance of the so-called D8 double-bubble aircraft concept proposed by a team led by the Massachusetts Institute of Technology (MIT), which incorporates enabling technologies in response to the NASA ‘N+3’ solicitation. With potential performance benefits of up to 70% decrease in per-passenger Specific Fuel Consumption (SFC) over a currently-flying baseline aircraft, the D8 aircraft design1, 2 is proposed as a Boeing 737 class aircraft that carries 180 passengers with a range of 3000 nautical miles (see Figure 1). The proposed aircraft has several advantages that allow it to meet the challenging performance metrics of NASA’s FW project. Among them are increased lift generated by the fuselage and a positive pitching moment at the cruise condition, resulting in a smaller wing and a smaller horizontal tail. These advantages were computationally investigated in a previous study.3 A key enabling technology in the D8 design is boundary layer ingestion into the propulsors. The nacelles are flush-mounted to the upper, rear surface of the double-bubble fuselage such that the fan ingests the boundary layer developing on the top of the fuselage. It was shown by Smith4 that there is a propulsive efficiency benefit gained from placing the fan in the wake or boundary layer of an aircraft, as commonly exploited in marine vehicles. Since part of the wake is being ingested by the propulsor, the entropy generation associated with the wake mixing is eliminated. This allows a reduction in propulsor shaft power, all else being equal. If this boundary layer ingestion can be done in a way that does not substantially increase dissipation on the airframe, we expect a 7-8% reduction in the power required for cruise flight based on control volume arguments,2 from aerodynamic effects alone. The focus of this paper is to demonstrate this benefit using viscous computations, validated with experimental results from a test in the NASA Langley 14x22 wind tunnel.5 Although the propulsive efficiency gain is the main benefit of the D8 design, there are several additional benefits that result from this design. First, most engine fans are designed to have a fan-face Mach number of M = 0.6, which is attained at a cruise speed of M = 0.72 by the rear D8 fuselage acting as a diffuser. This allows a minimal nacelle to be used for this BLI installation. The fuselage is also contoured to act as the lower part of the nacelles, while also serving as the mounting point for the vertical tails. These features result in a reduction of nacelle, size, weight, and drag compared to a podded or under-wing nacelle. To assess the benefit of BLI for the D8 airframe, a “podded” configuration representing an equivalent non-BLI aircraft is used as the comparison reference case. This podded configuration has the nacelles mounted on the sides of the rear fuselage, similar to a business jet. Computations are performed for both the integrated and the podded configurations and the results are compared to assess the BLI benefit. The metric of comparison is the mechanical flow power PK , defined by I ˆ ) dA . PK = (pt,∞ − pt ) (V · n (1) propulsor
PK is the mechanical power added to the flow by the propulsors. The comparison is performed at a given net stream-wise (horizontal) force on the whole aircraft. This PK is the power flow required to balance all the flow-field dissipation sources and follows from the power balance method developed by Drela.6 The integral in equation (1) is carried out over a control volume encompassing the propulsor, specifically the inflow and outflow planes. At the cruise point, the net horizontal force is zero (thrust equals drag), and PK is directly related to the engine shaft power required to maintain the cruise condition through PK = ηfan Pshaft .
(2)
Since the purpose of this study is to demonstrate the aerodynamic BLI benefit, we are concerning ourselves with PK rather than Pshaft and neglecting the impact of BLI on fan efficiency. A recent study7 has shown that the BLI fan efficiency penalty could be limited to 1%, but this result may vary for specific designs, and further research is needed. To compute PK , viscous computations are carried out using the structured overset mesh approach at wind tunnel (WT) conditions. The numerical modeling and solution methods are presented, followed by a discussion of the effects of various parts of the wind tunnel test model on the aerodynamics. Validation of the results based on the data from an unpowered 1:11 scale model of the D8 aircraft tested in the NASA Langley 14x22 WT are presented. It is shown that the BLI configuration requires 9% less mechanical flow power to maintain cruise condition (zero axial force at 2degrees angle of attack) than the podded configuration. This same benefit amount holds for a wide range of net stream-wise forces, corresponding to different flight 2 of 14 American Institute of Aeronautics and Astronautics
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(a) Unpowered
(b) Podded
(c) Integrated (BLI)
Figure 2. Configurations tested in the WT. Colors indicate overset surface patches.
angles. This stems from a fundamental propulsive efficiency improvement, rather than the details of the nacelle design.
II.
Experimental Setup
An experiment on a 1:11 scale model of the D8 aircraft was carried out in the NASA Langley 14x22 wind tunnel. This experiment is used to validate the CFD for the unpowered configuration. The experiment also examined the benefits of BLI. The details of the experiment are reported by Uranga et al.5 in a companion paper. A.
Model Configurations
Three versions of the D8 aircraft are used to explore various aspects of the design’s performance and are depicted in Figure 2. The simplest configuration, shown in Figure 2(a), is comprised of only the fuselage, wing, and tail surfaces, and is referred to as the “unpowered” configuration due to its lack of propulsors. We use the unpowered configuration to further validate the CFD methodology. In the “podded” configuration, shown in Figure 2(b), the nacelles are mounted at the rear of the fuselage, akin to a business jet or the MD-80. This is the baseline conventional (non-BLI) configuration for which the fan inlet flow is nearly uniform and at free-stream total pressure. The third “integrated” configuration, Figure 3. Photograph of the model in the NASA Lanshown in Figure 2(c), the engines are integrated into gley 14x22 wind tunnel. (George Homich, LARC). the top aft fuselage, between the two vertical fins, and ingest part of the fuselage boundary layer. The wing and forward-fuselage geometries are identical for all configurations, and the unpowered and podded configurations also have the same rear fuselage except for the engine pods. B.
Wind Tunnel Configuration
In the experiments at the NASA Langley 14x22 WT, a single-point support structure, shown in Figure 3, was used to mount the aircraft and alter the angle of attack. The modular design of the model allowed the experimentalists to remove the empennage (rear 20% of the fuselage with tail) and attach a different empennage to configure the various configurations. Boundary layer trips were employed to obtain turbulent flow on the wing, fuselage, tail surfaces, and nacelles.
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Figure 4. A cut through the overset mesh showing areas of mesh-to-mesh communication in blue.
The experiments were conducted at a nominal tunnel velocity of 70 mph, corresponding to Mach number of 0.083, and Reynolds number of 44440/in. Actual tunnel conditions are based on an empty-tunnel calibration using measurements from a set of pitot and static probes in the test-section entry cone.
III.
CFD Methodology
The Overflow code12 is used to obtain viscous solutions on the three configurations of the D8 aircraft with an actuator disk to model the fan. A methodology for structured overset mesh computations of the D8 airframe using Overflow was defined and validated in a previous study.3 Following that study and using the overset best practices,8 a baseline mesh is developed. The Chimera Grid Tools (CGT) package9 is used to generate surface and volume meshes.10 Because of symmetry, only the right half of the airplane and wind tunnel is modeled in all CFD simulations, thus roughly halving the computational cost compared to a full simulation. Overlapping surface grids are used to generate near-body volume meshes with a body-fitted grid covering the region near the WT walls. A set of Cartesian box grids cover the space between the near-body grids and the test section wall grid. Finally, the volumes upstream and downstream of the test section are covered with a core grid that follows the shape of the outer shell of the WT grid. Figure 4 shows a depiction of the overset meshes with blue regions indicating areas of mesh to mesh communication. A study of solution sensitivity to various meshing parameters was also carried out previously3 to determine the best parameters for a mesh. Four major parameters were tested independently: the wall spacing (target y + ), the surface spacing (wing leading edge spacing, trailing edge spacing, and global spacing parameters on the surface), the near-wall stretching ratio, and the off-body spacing. The study led us to use a y + of 1, a fine mesh reference value of 0.5, a stretching ratio of 1.07, and off-body spacing corresponding to 2.4% of chord is used. The surface spacing of 0.5 corresponds to a 0.016c spacing at the wing midspan. The leading edge spacing is 0.1% of chord, and the trailing edge spacing is half of the leading edge spacing. All other surfaces (e.g. fuselage, pi-tail) follow similar surface mesh spacing rules and are a function of the same reference value. These choices result in a surface mesh that has approximately 550 points defining the root airfoil and 300 points defining the tip airfoil with the mesh stretching to coarser spacing in flatter regions compared to the leading and trailing edges. The resulting baseline mesh on the unpowered configuration consists of 36 overset volume meshes containing approximately 113 million points for the D8, WT walls, and mounting strut. The nacelles and hub add another 13 grids and approximately 15 million points to the mesh for the podded nacelle configuration. The integrated nacelles, shown in Figure 4, have a more complex geometry that needs 28 additional grids which add 24 million grid points over the unpowered configuration. The internal X-rays module11 of the Overflow code is used to obtain mesh connectivity and interpolation coefficients to facilitate communication between overlapping grids. Best practices obtained from the previous study are also used for the flow solver. The right-hand side is
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(a) Contours of spurious excess stagnation pressure. Scalar dissipation is shown on top, matrix dissipation (0.25D) is shown on bottom.
1
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Figure 5. Comparison of dissipation schemes for the Overflow viscous simulations.
discretized using second-order central differencing. On the left-hand side, the Pulliam-Chausee diagonalized approximate-factorization scheme13 is used with a grid sequencing startup and multi-grid to lower the cost of converging the computations. The solver is run in steady-state mode with the low-Mach number preconditioner turned on and the SST turbulence model.14 Matrix dissipation was used to mitigate spurious total pressure errors. Its importance is indicated in Figure 5(a), which shows contours of spurious excess total pressure (Cpt > 0) on the symmetry plane of the fuselage. The scalar dissipation (top) generates a larger amount of spurious total pressure production than the matrix dissipation (bottom) option. To further quantify the effect of dissipation, the dissipation coefficients in the matrix scheme are reduced. Figure 5(b) shows velocity profiles from various dissipation schemes. D refers to the default matrix dissipation coefficients of 2.0 and 0.04 for second- and fourthorder dissipation respectively.15 The 0.5D and 0.25D refer to half and quarter of the default coefficients, respectively. From this study we can conclude that as the amount of artificial dissipation is decreased the velocity profile converges uniformly, with the matrix dissipation giving the smallest spurious total pressure errors. For this reason, the matrix dissipation option was used in all our computations. A typical solution convergence of the resulting solution process is shown in Figure 6(a). The L2-norm of the right-hand-side residual is plotted for all grids and it can be seen that the residual for all grids converge approximately 4 to 6 orders of magnitude. The L2-norm of the turbulence model residuals (not shown) converged between 2 and 6 orders of magnitude for all grids. The lift is also shown to converge to a single steady value in Fig. 6(b). Pressure contours on the surface of the unpowered configuration, along with Mach contours on the symmetry plane are shown in Fig. 6(c). A typical flow field around an aircraft with wakes behind the lifting surfaces (including the fuselage) is evident. A.
Modeling the Wind Tunnel
It has been shown3 that for computations of the D8 in MIT’s Wright Brothers Wind Tunnel, the tunnel walls and mounting hardware have a significant effect on the lift force, even when the assumption of inviscid tunnel walls is adequate. The LaRC 14x22 WT tests used a less intrusive support structure, which for simplicity was not modeled in the CFD calculations. Because this paper focuses on the relative BLI benefits between the different configurations, the effects of the mounting support, which is common for all configurations, is less important than if absolute results are the focus. Nevertheless, the effect of this new mounting hardware still needs to be investigated. All other aspects of the WT geometry including the settling chamber, contraction section, test section, and diffuser are modeled with slip walls (see Figure 7). The tunnel total-static pressure ratio and Reynolds number are specified to match the observed conditions at the pitot probe in the wind tunnel. Contours of stagnation pressure loss in Fig. 7(b) show that the stagnation pressure loss is zero everywhere except in the wake. To provide the basis for consistent comparisons, non-dimensionalized force and moment coefficients are obtained using the same reference chord and area as in the experiment.
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Residual History
Force/Moment History
Log10(L2 norm of RHS)
Total Lift Coefficient
-2
1 FuselageNose(1) L2 RHS FuselageBody(2) L2 RHS FuselageRear(3) L2 RHS FuselageTail(4) L2 0.8 RHS WingBodyCollar(5) L2 RHS WingInner(6) L2 RHS WingOuter(7) L2 RHS WingTip(8) L2 RHS0.6 htailTop(9) L2 RHS htail(10) L2 RHS hTailTip(11) L2 RHS VHCollar(12) L2 RHS 0.4 vertTailTop(13) L2 RHS vertTailBot(14) L2 RHS VFCollar1(15) L2 RHS wingwakebox(16) L2 0.2RHS vtailwakebox(17) L2 RHS htailwakebox(18) L2 RHS WTContraction(19) L2 RHS WTTest(20) L2 RHS 0 WTDiff1(21) L2 RHS 0 WTDiff2(22) L2 RHS fuselagebox(23) L2 RHS frontbox(24) L2 RHS wingbox1(25) L2 RHS wingbox2(26) L2 RHS wingbox3(27) L2 RHS empennagebox(28) L2 RHS rear1box(29) L2 RHS rear2box(30) L2 RHS rear3box(31) L2 RHS topbox(32) L2 RHS botbox(33) L2 RHS WTContrCore(34) L2 RHS WTDiff1Core(35) L2 RHS WTDiff2Core(36) L2 RHS Total Lift Coefficient
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(a) Residual. Downloaded by NASA AMES RESEARCH CENTER on January 21, 2014 | http://arc.aiaa.org | DOI: 10.2514/6.2014-0907
total CL Total
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(b) Lift history.
(c) Contours of Cp on the body surfaces, and Mach number in the symmetry plane Figure 6. A typical Overflow solution on the unpowered D8 aircraft in the LaRC 14x22 WT.
(a) Mesh.
(b) Contours of stagnation pressure loss coefficient for the unpowered configuration. Figure 7. The WT settling chamber, contraction section, test section, and diffuser.
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120 k
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(a) Podded.
(b) Integrated. Figure 8. Contours of stagnation pressure loss coefficient in a cutting plane going through the centerline of the propulsor.
B.
Modeling the Engine Fan
The propulsor fan was modeled as an actuator disk placed at the approximate fan location, and a uniform pressure rise was imposed across it. Since the fluid velocity is unchanged by the actuator disk, the pressure rise across the actuator disk is also the stagnation pressure rise. The computation is carried out with four different pressure rise values spanning the expected stagnation pressure rise at the cruise point. Figure 8 shows the stagnation pressure loss coefficient, (pt∞ − pt )/q∞ , in a cutting plane located approximately at the centerline of the propulsor for the podded and integrated configurations. The figure helps illustrate boundary layer ingestion in a visual way by illustrating the difference in the flow coming into the propulsor for the podded and integrated nacelles. Since the podded nacelle is behind the wing on the side of the rear fuselage, we expect it to be close to the wing wake. The wing wake is seen to be passing just under the nacelle and a uniform incoming stream is entering the nacelle as in a conventional engine mount on existing aircraft. In contrast to the podded configuration, we expect the fuselage boundary layer to be in the path of the inlet. Total pressure contours point to a thick boundary layer being ingested by the integrated nacelle. The boundary layer is seen to be growing rapidly in the diffuser region (aft fuselage) just ahead of the nacelle inlet. In both cases, the actuator disk has resulted in a total pressure rise across the fan face location and a plume is seen exiting the nacelles. Since the podded and integrated actuator disks are set at the same static pressure rise (and hence stagnation pressure rise), the main difference in propulsor performance comes from differences in mass flow. The non-dimensional propulsor mass flow m/(ρ ˙ ∞ V∞ Sref ) vs. non-dimensional pressure rise ∆pt /q∞ from the solutions for the two configurations can be seen in Figure 9. The integrated configuration has roughly 2–4% lower mass flow, largely due to the lower stagnation pressure in the ingested fuselage boundary layer.
IV.
Validation Against the Unpowered Configuration
A computational study of the external aerodynamic characteristics of the unpowered aircraft was presented previously,3 along with validation against an experiment done in the MIT Wright Brothers Wind Tunnel (WBWT) on a 1:20 scale model. To further validate the CFD methodology for the present config-
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Mass flow
urations, the computational and experimental results (lift, drag, and pitching moment coefficients) for the 1:11 scale unpowered configuration in the NASA Langley 14x22 WT are presented below for M = 0.088 and angles of attack between 0 and 8 degrees. Since our metric of comparison 0.04 is the mechanical flow power reIntegrated Podded quired at zero net force, it is important that we are able to accurately 0.035 predict the horizontal force (drag for the unpowered configuration). The drag comparison for the same 0.03 α sweep is shown in Fig. 10(b). The Overflow results match the WT data within the experimental un0.025 certainty of 8 drag counts, confirming that the CFD procedures used are accurate for prediction of the flow over the D8 aircraft, includ0.02 0 1 2 3 4 Stagnation pressure rise ing the performance metrics of interest: the axial force coefficient Figure 9. Propulsor non-dimensional mass flow (m/(ρ ˙ ∞ V∞ Sref )) vs. presCX = FX /(q∞ Sref ), and the me- sure rise (∆pt /q∞ ). chanical power coefficient CPK = PK /(q∞ V∞ Sref ).
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Figure 10. Comparison of force and moment coefficients (14x22 WT data vs. Overflow2 viscous CFD) on the unpowered configuration. The vehicle is modeled with wind tunnel walls and without mounting hardware.
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Figure 10(a) shows the lift coefficient as a function of angle of attack. The lift coefficients from the Overflow results compare well to the experimental results with a slightly higher lift prediction than the wind tunnel values at higher angles of attack. This was also observed in the previous study3 and it was shown that the turbulence model has a large effect on the lift at high angles of attack. The pitching moment is also shown in Figure 10(c) and is in reasonable agreement with the experiment.
Results
The overall benefit of the BLI configuration over the podded configuration is demonstrated in Figure 11, where the mechanical flow power coefficient CPK = PK /(q∞ V∞ Sref ) is plotted against the net force coefficient, CX = FX /(q∞ Sref ) for both podded and integrated configurations for several values of the actuator disk pressure jump. As previously stated, this is a direct comparison of how much power the propulsors need to put into the flow to maintain the cruise point requirement CX = 0. CX is computed by taking the axial component of the integrated pressure and viscous forces on the airframe solid surfaces and also on the actuator disk. CPK is computed from its definition (1), with the integral evaluated on each side of the actuator disk. As expected from control volume analysis, the integrated configuration requires 9% less mechanical flow power to maintain CX = 0 (the cruise point) than the podded (CPK = 0.04159 vs. 0.04583). This difference of 0.004 in CPK is approximately maintained throughout the range of CX examined. The LaRC 14x22 WT experiments measured a BLI benefit of 6% less electrical power required for the integrated configuration. This is a valid metric for comparison only if the efficiencies of the fan and electric motor are unchanged between podded and integrated configurations. Additional experiments are underway to relate the experimentally measured electrical power to PK . To determine the mechanisms that produce the difference in performance, we examine cut planes for each configuration (shown in Figure 12) at three locations: one at the integrated nacelle inlet, one at the fuselage trailing edge, and one in the wake 5 inches downstream of the integrated nacelle trailing edge. Contours of stagnation pressure loss on the inlet cutting planes are plotted in Figure 13 for all three configurations. The total pressure defects in the boundary layer and the wing wake can be seen in all three configurations. A comparison of the podded and unpowered configurations indicates that the presence of the pods does not 0.15
Integrated (CFD) Podded (CFD)
0.1
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V.
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CX Figure 11. Non-dimensionalized required mechanical flow power coefficient CPK to maintain specified level of net force coefficient CX for both podded and integrated configurations. At CX = 0 (interpolated), the integrated (BLI) configuration requires 9% less mechanical flow power than the podded (non-BLI).
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Figure 12. Schematic of the cut locations in front of the inlet and behind the aircraft.
(a) Unpowered
(b) Podded
(c) Integrated
Figure 13. Contours of stagnation pressure loss coefficient at the integrated nacelle inlet plane for unpowered, podded, and integrated configurations (X = 108").
affect the fuselage boundary layer, which was the design intent. We quantify the entropy generation upstream of the cut plane to show that the re-distribution does not generate excess loss. From power balance,6 we have the dissipation in a given control volume, Φ = E˙ − PK ,
(3)
where PK is the mechanical energy flow rate into a control volume from the propulsor defined by equation (1), and E˙ is the mechanical energy flow rate convected out of the control volume defined as I ˆ ) dA . E˙ = (pt∞ − pt ) (V · n (4) CV
This is the same as equation (1), but equation (4) is integrated over the control volume surface while equation (1) is integrated over the actuator disk. We are interested in the dissipation upstream of the cut plane and so we envision a control volume in which the tunnel walls are the outer surfaces, the upstream surface is the test section inlet, and the downstream surface is the cut plane. E˙ is non-zero only on the cut-plane. So for an actuator disk located upstream of the plane, the dissipation in the control volume is the integral of equation (4) on the cut-plane minus PK . We separate out the entropy generation from the different components, fuselage, wing, and podded nacelle, by integrating over specific non-overlapping areas of the plane, as illustrated in Figure 14. The results at the nominal power setting are shown in Table 1. Variations in the fuselage and wing dissipations for both podded and integrated configurations are very small compared to the unpowered. The presence of the podded nacelle provides the only major difference from the unpowered configuration at the inlet location. 10 of 14 American Institute of Aeronautics and Astronautics
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The boundary layer ingestion fraction fBLI is the fraction of Fuselage Region the kinetic energy thickness of the flow entering the nacelle inlet to the fuselage kinetic energy thickPodded Nacelle Region Wing Wake Region ness and is an important parameter determining the BLI benefit. The computed fBLI is 0.394, which is very close to the design value of 0.4. Next, we examine the wake cut behind the aircraft (X=130"). The wake signature of the wing, the empennage, and the fuselage can be seen in Figure 15. The wake for Figure 14. Schematic of the division of the integration plane to different components. the unpowered case, shown in Figure 15(a), indicates the highest total pressure loss occurring behind the fuselage/empennage junction and in the wake of the fuselage. For the podded case in Figure 15(b) the total pressure loss field in the wake is nearly identical to the unpowered case, except of course for the large total pressure surplus in the jet and a slight difference in the wake of the pylon and nacelle. The wake of the integrated configuration, shown in Figure 15(c), is similar to the podded in the wing and tail regions, but naturally differs behind the fuselage since much of the fuselage wake fluid has been ingested and now appears in the fan stream. The jet and wake are already in the process of mixing, and so further information could be obtained from a cut further upstream before mixing occurs. Table 1. Entropy generation coefficient ζ = Φ/q∞ V∞ Sref upstream of nacelle inlet plane.
Total
Fuselage
Wing
Nacelle
Unpowered Podded Integrated
0.02148 0.02425 0.02241
0.01067 (50%) 0.01071 (44%) 0.01129 (50%)
0.01080 (50%) 0.01132 (47%) 0.01112 (50%)
0 0.0022 (1%) 0
Variation
0.00228
0.00062
0.00052
0.0022
A more useful cut location is the fuselage trailing edge (X=117"), since the effect of the fuselage surface dissipation is complete, but the wake has not had time to mix yet. The contours of stagnation pressure loss at this cut location are shown in Figure 16 for all three configurations. The fuselage wake structure at this location is much more clearly defined. The integrated fuselage wake in Figure 16(c) is smaller in area, since a portion has been ingested. However, there are also regions between the engines and at the nacelle-tail junction in which the stagnation pressure (Cpt ≈ 1) is much lower than that in the podded fuselage wake in Figure 16(b). We may conclude from this that there are regions in which the flow has separated and that additional performance improvements in the integrated configuration may be obtained by eliminating this separation through a suitable redesign. At the X=117" cut plane, the only remaining surface dissipation is on the pi-tail, and this is not expected to differ between configurations. Thus the only remaining entropy generation of interest occurs in the mixing of the jet and wake with the freestream. This can be estimated via equation (5), in which it is assumed that the flow expands isentropically to the freestream pressures and mixes with the freestream flow. Z 1 ˆ dA T ds = (Vexp − V∞ )2 ρV · n (5) 2 The entropy generation is divided into four components: the jet mixing, the fuselage wake, the wing wake, and the nacelle. For the integrated configuration, the nacelle and fuselage are analyzed together. The jet mixing loss is computed directly from the integral of equation (5) on the downstream face of the actuator disk, and this is regarded as an airframe performance parameter. The upstream dissipation in the fuselage wake, wing wake, and nacelle are computed as in Figure 14. The mixing loss computed for each region is added to this, and the jet mixing is deducted from the nacelle component. 11 of 14 American Institute of Aeronautics and Astronautics
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(a) Unpowered
(b) Podded
(c) Integrated
(d) Unpowered
(e) Podded
(f) Integrated
Figure 15. Contours of stagnation pressure loss coefficient at the wake cut plane (X=130") for unpowered, podded, integrated configurations.
The results are presented in Table 2. Comparing the podded and unpowered configurations, it is clear from both the integrated quantities and the contours that the podded propulsor does not affect the airframe since the fuselage and wing dissipation are the same for each. The podded nacelle with the jet contributes an additional dissipation of ζ = 0.101 compared to the unpowered configuration. The comparison between integrated and podded configurations is more interesting. The integrated configuration has a 6% lower overall dissipation than the podded configuration. Three quarters of this benefit comes from the lower jet velocity, and one quarter comes from a slight decrease in fuselage/nacelle dissipation, despite the regions of separated flow noted above. Table 2. Entropy generation coefficient ζ = Φ/(q∞ V∞ Sref ) in wake plane.
Unpowered Podded Integrated
Total
ζfuse
ζnacelle
ζjet
ζwing
0.0240 0.0345 0.0325
0.0129 (54%) 0.0129 (38%) 0.0149 (46%)
0 0.0025 (7%) 0
0 0.00760 (22%) 0.00603 (19%)
0.0111 (46%) 0.0115 (33%) 0.0115 (35%)
The integrated configuration has lower jet velocity than the podded because the integrated propulsor inlet flow has lower stagnation pressure than freestream. Thus, for the same pressure rise, the integrated exit flow has lower stagnation pressure than the podded configuration. To show this quantitatively, the mass-averaged jet stagnation pressure (pt,jet )avg /q∞ is plotted against the net axial force (CX ) in Figure 17(a). The integrated configuration mass flow is also about 2% lower (Figure 17(b)) for the same CX . Lower stagnation pressure and lower mass flow lead to lower jet dissipation for the integrated configuration across the range of CX examined and this is shown in Figure 17(c).
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(b) Podded
(c) Integrated
Figure 16. Contours of stagnation pressure loss coefficient at the fuselage trailing edge plane for unpowered, podded, and integrated configurations (X = 117”). 4
0.04
Integrated Podded
3
1
0.03
-0.03
-0.02
-0.01
0
0.01
0.02
0.03
0.04
0.02 -0.05
CX
0.02
0.01
0.025
-0.04
-0.04
-0.03
-0.02
-0.01
0
0.01
0.02
0.03
0.04
CX
(a) Jet Stagnation Pressure
Integrated Podded
0.03
Φjet/V∞q∞Aref
2
0 -0.05
0.04
Integrated Podded
0.035
Mass flow
(Pt,jet-P∞)/Q∞
Downloaded by NASA AMES RESEARCH CENTER on January 21, 2014 | http://arc.aiaa.org | DOI: 10.2514/6.2014-0907
(a) Unpowered
(b) Jet Mass Flow
0 -0.05
-0.04
-0.03
-0.02
-0.01 CX
0
0.01
0.02
0.03
(c) Jet Dissipation
Figure 17. Jet stagnation pressure (pt,jet,M A − p∞ )/q∞ , non-dimensional mass flow m/(ρ ˙ ∞ V∞ Aref ), and dissipation Φ/(q∞ V∞ Aref ), plotted against net axial force coefficient CX for podded, integrated configurations.
VI.
Concluding Remarks
The aerodynamic benefit of boundary layer ingestion (BLI) has been demonstrated by viscous overset computations of D8 configurations with BLI and non-BLI propulsors. The overset computational capability is validated against test data from NASA Langley 14x22 Wind Tunnel tests for a 1:11 scale unpowered model. A comparison of the mechanical flow power computed from the simulation results for a conventional podded nacelle configuration and an integrated nacelle configuration shows that the BLI benefit for this particular BLI configuration of the D8 is 9%, despite losses from flow separation in the integrated configuration empennage. Additional examination of the inlet and wake flow is conducted to characterize the mechanisms that produce this difference in performance. This investigation shows that the integrated configuration results in 6% lower overall dissipation compared to the podded configuration. It also shows that the integrated configuration has lower stagnation pressure and lower mass flow compared to the podded configuration which results in lower jet dissipation for all values of axial force. This result indicates that BLI has the potential to reduce fuel burn rate for commercial aircraft and therefore merits further study. To maximize the benefits of BLI for the D8 configuration, the computations shown will be used to iterate on the integrated configuration design to regain performance losses (increases in dissipation) relative to the unpowered configuration. Care also needs to be taken to avoid excessive dissipation at different operating conditions (takeoff, high α, etc...). Additionally, the uniform pressure jump actuator disk model that was used in these computations is not entirely representative of what happens when the fan processes an inlet distortion. In a real fan, the lower stagnation pressure fluid experiences a larger pressure rise than the high stagnation pressure fluid and this has implications for the flow behavior upstream of the inlet as well as the character of the nozzle exhaust flow. Future computations are planned with advanced actuator disks implemented to determine the importance of this phenomenon for BLI configurations. Finally,
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this work needs to be extended to the Mach and Reynolds numbers of a full-scale configuration in free flight.
Acknowledgements Support for this work is provided by the Fixed Wing Project of the NASA Fundamental Aeronautics Program. NASA Advanced Supercomputing division provided the Computing resources used for the project. The authors wish to thank Dr. Mark Drela and Dr. Ed Greitzer for fruitful discussions and suggestions. We also wish to thank Dr. William Chan for mesh generation advice and to Dr. Pieter Buning for advice on flow solver options and boundary conditions.
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2 Greitzer,
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