8th Annual Shock Wave/Boundary Layer Interaction Technical Interchange Meeting14 April, 2015

Inward-Turning Inlets for Low-Boom / Low-Drag Applications

Chuck TrefnyJohn SlaterSam Otto

NASA Glenn Research Center

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)

Axisymmetric Spike Two-DimensionalAxisymmetric Pitot STEX-Circular STEX-Flattop

Encouraging preliminary results, but distortion levels were higher than desired...

Supersonic compression surface is generated by tracing streamlines through an inward-turning “parent” flowfield

“Inward-Turning” Inlets

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

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

“Inward-Turning” Inlets

Streamline-traced shape from circular tracing curve

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

Truncation of the Busemann Flowfield is Required

”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...

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

“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 1Match Mach number and flow deflection angle

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

Merging Approach 3Match ICFA expanded Mach number and ray angle

Merging Approach 3Final Design – Reduce Exit Mach Number

Streamline-Traced Contour from Merging Approach 3

Traced from circular throat, tangent to parent flowfield axis

Focal Point

Parent Flowfield Axis

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

“Vent Region” Modification“Native” Geometry

Subsonic Diffuser and Nozzle Added for RANS Simulation

Summary of Inlet Performance Based on RANS Solutions

RANS Simulation of Back-Pressure Effect – No Bleed

a

d

c

b

Bleed Simulation in RANS Solutions

RANS Simulation of Back-Pressure Effect ~2% Bleed

a

d

c

b

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...

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

Simple Truncation Results in Non-Uniform Flow

Design Space – Recovery vs. Length

Design Space – Recovery vs. Outflow Mach

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

Opportunity to Leverage Aerion Test

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