Inward-Turning Inlets for Low-Boom / Low-Drag Applications Chuck Trefny John Slater Sam Otto NASA Glenn Research Center 8th Annual Shock Wave/Boundary Layer Interaction Technical Interchange Meeting 14 April, 2015

Outline Motivation Inward-Turning Inlets and Streamline Tracing Slater’s “STEX” Inlets New Flowfield Architecture and Otto’s Merging Procedure Mach 1.7 Inlet Design and Preliminary CFD Results Proposed 8x6 Test

Slater Introduces Inward-Turning “STEX” Inlets (2014) Streamline-traced, external-compression, or “STEX” inlets introduced to the commercial supersonic community as a way to reduce sonic boom and cowl drag at performance levels comparable to traditional inlet types Really are internal compression with a twist... Axisymmetric Pitot Axisymmetric Spike Two-Dimensional STEX-Circular STEX-Flattop Encouraging preliminary results, but distortion levels were higher than desired...

“Inward-Turning” Inlets Supersonic compression surface is generated by tracing streamlines through an inward-turning “parent” flowfield Axisymmetric “Busemann” flowfield normally used as the parent flowfield – short and efficient This is a 100% internal compression inlet Conical, isentropic compression where conditions are constant along rays that emanate from a focal point. Described by the Taylor-Maccoll equations Non-circular tracing plane shapes may be used to generate a conformal inlet aperture Busemann (1942) “conical” compression

“Inward-Turning” Inlets Supersonic compression surface is generated by tracing streamlines through an inward-turning “parent” flowfield Streamlines are traced from an arbitrary “tracing curve” in a plane at the compression field exit, forward to freestream conditions Axisymmetric “Busemann” flowfield normally used as the parent flowfield – short and efficient This is a 100% internal compression inlet Conical, isentropic compression where conditions are constant along rays that emanate from a focal point. Described by the Taylor-Maccoll equations Non-circular tracing plane shapes may be used to generate a conformal inlet aperture

“Inward-Turning” Inlets Supersonic compression surface is generated by tracing streamlines through an inward-turning “parent” flowfield Streamlines are traced from an arbitrary “tracing curve” in a plane at the compression field exit, forward to freestream conditions The resulting inlets are inherently “internal compression” and would exhibit nonlinear “start/unstart” flow phenomena Off-axis placement of the tracing curve mitigates “starting” issue Axisymmetric “Busemann” flowfield normally used as the parent flowfield – short and efficient This is a 100% internal compression inlet Conical, isentropic compression where conditions are constant along rays that emanate from a focal point. Described by the Taylor-Maccoll equations Non-circular tracing plane shapes may be used to generate a conformal inlet aperture Streamline-traced shape from circular tracing curve

Truncation of the Busemann Flowfield is Required Leading ray of Busemann compression is a Mach wave at freestream conditions and zero deflection angle Length of full Busemann flowfield is prohibitive, many truncation studies in the literature for the hypersonic application For the low-boom application, initial inward deflection is required to reduce or eliminate the external nacelle angle, drag, and boom Structural thickness results in frontal area if not truncated

”STEX” Inlet Design Procedure Initial cowl angle imposed, and blended into stream-traced contour Terminal shock forced by back-pressure Uniform, isentropic properties of parent flowfield compromised New parent flowfield architecture was proposed... Modified design procedure is proposed to improve recovery and distortion...

New Parent Flowfield Architecture Include leading oblique wave in parent flowfield Terminal shock also included in parent flowfield by using “strong” oblique wave as Busemann exit shock

Internal Conical Flow A (Molder, 1967) Solution to the Taylor-Maccoll equations marching downstream from oblique wave to a singular point Conditions on the singular ray must be merged with the truncated Busemann flowfield Key is to merge conditions on the singular ray to those on a ray of the Busemann flowfield “ICFA” flowfield nomenclature

Merging of ICFA and Busemann Flowfields Mach number, ray angle, and flow deflection angle on the ICFA singular ray cannot all be matched to a Busemann truncation ray Flow non-uniformity depends on approach to merging...

Merging Approach 1 Match Mach number and flow deflection angle

Merging Approach 2 (You, et al., 2009) Match Mach number and ray angle

Merging Approach 3 Match ICFA expanded Mach number and ray angle

Merging Approach 3 Final Design – Reduce Exit Mach Number

Streamline-Traced Contour from Merging Approach 3 Traced from circular throat, tangent to parent flowfield axis Parent Flowfield Axis Focal Point

Modifications to the Native Geometry for Viscous Effects Compression surface displaced outward to accommodate boundary-layer displacement thickness “Shoulder” rounded to ease shock interaction and provide better off-design performance “Vent Region” modified to facilitate starting and sub-critical spillage “Native” Geometry “Vent Region” Modification

Subsonic Diffuser and Nozzle Added for RANS Simulation

Summary of Inlet Performance Based on RANS Solutions Parent flowfield at 0.989 recovery, so 7% loss due to imperfect merging and viscous losses for no-bleed case. Recovery increased to Mil Std with roughly 2% bleed. GE Method “D” distortion values all near the edge of the F404 operating envelope In general, distortion is reduced with back-pressure and bleed

RANS Simulation of Back-Pressure Effect – No Bleed Parent flowfield is intact at critical point “b” Momentum deficit at 12 o-clock is obvious Point out SWBLI b a

Bleed Simulation in RANS Solutions

RANS Simulation of Back-Pressure Effect ~2% Bleed Parent flowfield is nicely preserved, and distortion is reduced Point out non-linearity b a

Summary New design scheme for inward-turning, low-boom inlets developed with leading shock included in parent flowfield, and “strong” terminal oblique wave Analytical merging of ICFA and Busemann flows validated by Euler analysis Mach 1.7 design validated with 3-D Turbulent RANS Roughly 2% boundary-layer bleed improved recovery to MIL-E-5007D Non-linearity noted in sub-critical characteristics in bleed case 8x6 test proposed for experimental validation

Objectives of Proposed 8x6 Wind-Tunnel Testing Validate the effects of bleed and other boundary-layer control schemes such as vortex generators on overall inlet performance Better understanding of non-linear sub-critical phenomena Determine tolerance to angles of attack and yaw Determine off-design Mach number performance

Back-Up

STEX Inlet Design Procedure Initial cowl angle imposed, and blended into stream-traced contour Terminal shock forced by back-pressure Uniform, isentropic properties of parent flowfield compromised Modified design procedure is proposed to improve recovery and distortion...

Simple Truncation Results in Non-Uniform Flow Initial deflection results in a curved shock wave and Mach disk at the parent flowfield axis Conditions downstream of the non-isentropic shock wave cannot match those of the conical flow on any ray Parent flowfield is compromised resulting in total pressure loss and non-uniform flow at the exit

Design Space – Recovery vs. Length

Design Space – Recovery vs. Outflow Mach

Opportunity to Leverage Aerion Test Reimbursable program in late FY15 NASA is modifying existing adapter to include AIP instrumentation Cold-pipe and mass flow plug are existing Opportunity to test NASA configuration as follow-on NASA Inlet Adapter Cold-Pipe Mass Flow Plug

Schedule and Budget Two-week test begins 12 mo. from go-ahead Final report 18 mo. from go-ahead ROM cost for fab and test based on similar, recent 8x6 tests: $1.5M