1. 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

2. 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

3. 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...

4. “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

5. “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

6. “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

7. 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

8. ”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...

9. 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

10. 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

11. 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...

12. Merging Approach 1
Match Mach number and flow deflection angle

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

14. Merging Approach 3
Match ICFA expanded Mach number and ray angle

15. Merging Approach 3
Final Design – Reduce Exit Mach Number

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

17. 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

18. Subsonic Diffuser and Nozzle Added for RANS Simulation

19. Summary of Inlet Performance Based on RANS Solutions
Parent flowfield at 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

20. 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

21. Bleed Simulation in RANS Solutions

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

23. 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

24. 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

25. Back-Up

26. 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...

27. 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

28. Design Space – Recovery vs. Length

29. Design Space – Recovery vs. Outflow Mach

30. 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

31. 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