Notes

Outline

Jet Engine Inlet Design Cheryll Hawthorne Supervised by Dr. Alvi EML 4421 09 Nov 01

Topics Subsonic Inlets Flow Patterns Internal Flow External Flow Inlet Performance Criteria Supersonic Inlets Reverse Nozzle Diffuser Shock Boundary Layer Problem External Deceleration Flow Stability Problem

Design Objectives Prevent boundary layer separation Lower sensitivity to pitch and yaw Minimize stagnation pressure loss Produce uniform flow velocity and direction Increase efficiency operation in both supersonic and subsonic Reduce flow distortion at engine fan face Increase pressure recovery

Jet Engine Components Inlet-sucks in air Compressor-squeezes the air Combustor-adds heat to the air Turbine-provides work for the squeezing process Nozzle-blows the air out the back

Engine Layout

Inlet Sucks in air Slows air down Feeds air into compressor and  fans

Inlet Air Flow Subsonic Supersonic-use shock wave to slow down air

Air-Breathing Engines Based on Gas generator

Types of Air-Breathing Engines Turbojet Turbofan Afterburning Turbofan Turboprop/ shaft Ramjet Scramjet Turbojet/Ramjet

Various Inlet Models Ramjet Scramjet Turbojet/Ramjet Combo

Ramjet Ramjet Incoming high speed air Compressed by ram effect For high enough air speed, no compressor or turbine needed

Scramjet Scramjet Supersonic Combustion Ramjet Air mixed with fuel while traveling at supersonic speeds Temp increase and pressure loss due to shocks are greatly reduced

Pulse Jets Pulse Jets Series of spring-loaded shutter type valves before compressor Valves close to prevent backflow

Background & Motivation Pressure and/or velocity flow distortions at engine (compressor) fanface can compromise engine efficiency. Separation of incoming boundary-layer flow can reduce pressure recovery and lead to: Unsteady loading Increased fatigue of engine fan blades Aerodynamic stall on compressor blades1

Integrated Propulsion Systems Joint Strike Fighter NASA/Boeing, Blended Wing Body

Boeing JSF X-32B Joint Strike Fighter

Blended Wing Body Engine inlets located at the aft end of  aircraft Developing large boundary layer upstream of engine inlet

YB-49 Northrop Blended Wing Body

Subsonic Inlets Inlet operates with a wide range of incident stream conditions due to flight speed and the mass flow demand by the engine

Inlet Area chosen to minimize external acceleration during takeoff Upstream area is less than inlet area

Compressor Inlet Conditions Stagnation Temperature T02=Ta(1+M2(k-1)/2) Stagnation Pressure P02=pa(1+nd(T02/Ta)-1))kd/(kd-1) nd=adiabatic diffuser efficiency

Inlet Flow Behaves as though in a diffuser Momentum decreases Pressure rises No work

Flow Patterns Inlet area often chosen to minimize external acceleration during takeoff So that external deceleration occurs during level-cruise operation External deceleration requires less internal pressure rise Hence, less severe loading of the boundary layer

Internal Flow Flow in the inlet behaves like a diffuser or decelerator Inlet design depends on: Potential flow calculations Boundary layer calculations Wind tunnel testing to assess inlet performance under a wide range of test conditions

Separation in the Inlet Separation may take place in 3 zones External flow zone Along underside of internal flow zone Along upperside of lower wall of internal flow zone At high angles of attach, all three zones could be subjected to unusual pressure gradients

External Flow Inlet design requires a compromise between external and internal deceleration to prevent boundary layer separation

Boundary Layer Separation in Subsonic Flow Subsonic flow over inlet lip High velocity causes low pressure region followed by high pressure region Causing boundary layer separation

Boundary Layer Separation in Supersonic Flow Supersonic flow usually ends in abrupt shock Shock wall intersection may cause boundary layer separation

Shock-Boundary Layer Problem For strong shock wave M>1.25 Large pressure gradient near wal Fluid near wall cannot move in main direction Boundary layer separates

Boundary Layer Separation must be Avoided Results in poor pressure recovery in the flow Causing extra rearward drag on the body Decreasing efficiency

What is a Boundary Layer Boundary layers  separate from a body due to increasing fluid pressure in the direction of the flow (adverse pressure gradient) Increase in the fluid pressure increases potential energy of the fluid  kinetic energy decreases Fluid slows and boundary layer thickens Wall stress decreases and fluid no longer adheres to the wall

Boundary Layer Velocity Profile 2Boundary layers occur on surface of bodies in viscous flow

Laminar Boundary Layer

Viscosity causes boundary layer separation

Consequences of Boundary Layer Separation large increase in drag on the body Flow distortions

Passive Boundary Layer Control Methods Passive Uses vortex generators Supersonic microjets Enhance flow uniformity Boundary layer fluid is energized

Drawbacks to Passive Control Methods Drawback Performance is not uniform over entire engine Possible Solution Use large number of generators in inlet ducts Consequence Additional pressure loss

Active Control Methods active flow control scheme with feedback control Leads to reduced distortion over large parametric range

Separation may occur…. In zone 1 due to local high velocities and deceleration over outer surface In zone 2 or zone 3 depending on the geometry of the duct and the operating conditions

Inlet Performance Depends on the pressure gradient on both internal and external surfaces External pressure rise is fixed by:  external compression Ratio of Area Max   Area Inlet Internal pressure rise depends on the reduction of velocity between entry to the inlet diffuser and entry to compressor

Inlet Performance Criteria Isentropic Efficiency Stagnation pressure ratio

Isentropic Efficiency

Stagnation Pressure Ratio

Supersonic Inlets Flow leaving inlet system must be subonic Fully supersonic stream would cause excessive shock losses in compressor Mach number for flow approaching subsonic compressor:  Mmax=0.4-0.6

Mach Number Limits 4<M<6 approaching a subsonic compressor

RAMJET No Mach # limitations for RAMJET SCRAMJET – supersonic combustion ramjet However, no application to date in flight vehicle Causes excessive aerodynamic loss

Supersonic Inlets The Starting Problem The Shock-Boundary Layer Problem Flow Stability Problem

The Starting Problem Internal supersonic deceleration in a converging passage of nonporous walls is hard to establish Current solution-overspeeding the inlet air or varying the diffuser geometry

The Shock-Boundary Layer Problem Wall boundary layer may cause strong shocks A disastrous effect on duct flow Large shocks may require 10 duct widths or more to return to uniform flow

Current solutions Oblique shock - produces less pressure rise Create shock near thinnest part of boundary layer

Flow Stability Problem Subcritical-spilling of flow and normal shock upstream of inlet Critical-differs only in the amount of spillage Supercritical-normal shock occurs at a higher Mach #

Supersonic Diffusers Different geometries under testing However, diverters create additional drag

Other Considerations Shorten inlet lengths-reduce flow separation Vortex generators-energize boundary layer

Passive Boundary Layer Control Devices Reduce flow distortion by redistributing energy But performance of control devices not uniform over entire area Need large number of devices to achieve uniform performance

Proposed Active Boundary-Layer Control Scheme Use supersonic microjets to reduce distortion over large parametric range Grid of supersonic microjets installed in ramp Microjets placed at curve of ramp where separation is assumed

Monitor Flow Control Mean and unsteady surface flow properties are monitored near boundary layer separation Unsteady surface pressures measured with high frequency miiature pressure transducers Visualization techniques

Analysis Mean, total pressure contours obtained in cross planes at selected streamwise locations Contours represent effect of microjets on  steady-state distortion and total pressure recovery Measure pressure fluctuations above ramp to characterize dynamic distortion

Initial Tests Subsonic wind tunnel Initial tests will later be used to develop supersonic tests

References  Active Control of Boundary-Layer Separation & Flow Distortation in adverse Pressure Gradient Flows via Supersonic Microjets, proposal to NASA Langley Research Center, Farrukh Alvi http://www.desktopaero.com/appliedaero/blayers/blayers.html http://www.aircraftenginedesign.com/abefs.html Alvi, Elavarasan, Shih, Garg, and Krothapalli, “Active control of Supersonic Impingin Jets using Micro Jets, AIAA 2000-2236, submitted to AIAA Journal

Calculate the diffuser efficiency in terms of the Mach Number