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MAE 5350: Gas Turbines. Subsonic and Supersonic Inlets Mechanical and Aerospace Engineering Department Florida Institute of Technology D. R. Kirk. OVERVIEW: INLETS AND DIFFUSERS. Purpose: Capture incoming stream tube (mass flow)
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MAE 5350: Gas Turbines Subsonic and Supersonic Inlets Mechanical and Aerospace Engineering Department Florida Institute of Technology D. R. Kirk
OVERVIEW: INLETS AND DIFFUSERS • Purpose: • Capture incoming stream tube (mass flow) • Condition flow for entrance into compressor (and/or fan) over full flight range • At take-off (M0~0), accelerate flow to 0.4 < M2 < 0.7 • At cruise (M0~0.85), slow down flow to 0.4 < M2 < 0.7 • Remain as insensitive as possible to angle of attack, cross-flow, etc. • Requirements • Bring inlet flow to engine with high possible stagnation pressure • Measured by inlet pressure recovery, pd = Pt2/Pt1 • Provide required engine mass flow • May be limited by choking of inlet • Provide compressor (and/or fan) with uniform flow
EFFECT OF MASS FLOW ON THRUST VARIATION • Mass flow into compressor = mass flow entering engine • Re-write to eliminate density and velocity • Connect to stagnation conditions at station 2 • Connect to ambient conditions • Resulting expression for thrust • Shows dependence on atmospheric pressure and cross-sectional area at compressor or fan entrance • Valid for any gas turbine
NON-DIMENSIONAL THRUST FOR A2 AND P0 • Thrust at fixed altitude is nearly constant up to Mach 1 • Thrust then increases rapidly, need A2 to get smaller – WHY mathematically/physically?
CONTROL VOLUME ANALYSIS: SECTION 6.2 Aerodynamic force is always favorable for thrust production
OPERATIONAL OVERVIEW High Thrust for take-off Low Speed, M0 ~ 0 High Mass Flow Stream Tube Accelerates Lower Thrust for cruise High Speed, M0 ~ 0.8 Low Mass Flow Stream Tube Decelerates Aerodynamic force is always favorable for thrust production
OPERATIONAL OVERVIEW Low Thrust High Speed Low Mass Flow Stream Tube Decelerates High Thrust Low Speed High Mass Flow Stream Tube Accelerates
INLETS OVERVIEW: SUPERSONIC INLETS • At supersonic cruise, large pressure and temperature rise within inlet • Compressor (and burner) still requires subsonic conditions • For best hthermal, desire as reversible (isentropic) inlet as possible • Some losses are inevitable
SUPERSONIC INLETS Normal Shock Diffuser Oblique Shock Diffuser
NORMAL SHOCK TOTAL PRESSURE LOSSES Example: Supersonic Propulsion System • Engine thrust increases with higher incoming total pressure which enables higher pressure increase across compressor • Modern compressors desire entrance Mach numbers of around 0.5 to 0.8, so flow must be decelerated from supersonic flight speed • Process is accomplished much more efficiently (less total pressure loss) by using series of multiple oblique shocks, rather than a single normal shock wave • As M1 ↑ p02/p01 ↓ very rapidly • Total pressure is indicator of how much useful work can be done by a flow • Higher p0→ more useful work extracted from flow • Loss of total pressure are measure of efficiency of flow process
REPRESENTATIVE VALUES OF INLET/DIFFUSER STAGNATION PRESSURE RECOVERY AS A FUNCTION OF FLIGHT MACH NUMBER
C-D NOZZLE IN REVERSE OPERATION (AS A DIFFUSER) Not a practical approach!
MAE 5350: Gas Turbines Gas Turbine Engine Nozzles Mechanical and Aerospace Engineering Department Florida Institute of Technology D. R. Kirk
OVERVIEW: NOZZLES • Subsonic Aircraft: Usually a fixed area convergent nozzle is adequate • Can be more complex for noise suppression • Supersonic Aircraft: More complex, variable-area, convergent-divergent device • Two Primary Functions: • Provide required throat area to match mass flow and exit conditions • Efficiently expand high pressure, high temperature gases to atmospheric pressure (convert thermal energy → kinetic energy)
REVIEW: OPERATION OF CONVERGING-DIVERGING NOZZLES All practical aerospace nozzles operate in regimes (e)-(g)
REVIEW: OPERATION OF C-D NOZZLES • Figure (a) shows the flow through the nozzle when it is completely subsonic (i.e. nozzle isn't choked). The flow accelerates out of the chamber through the converging section, reaching its maximum (subsonic) speed at the throat. The flow then decelerates through the diverging section and exhausts into the ambient as a subsonic jet. Lowering the back pressure in this state increases the flow speed everywhere in the nozzle. • Further lowering pb results in figure (b). The flow pattern is exactly the same as in subsonic flow, except that the flow speed at the throat has just reached Mach 1. Flow through the nozzle is now choked since further reductions in the back pressure can't move the point of M=1 away from the throat. However, the flow pattern in the diverging section does change as the back pressure is lowered further. • As pb is lowered below that needed to just choke the flow a region of supersonic flow forms just downstream of the throat. Unlike a subsonic flow, the supersonic flow accelerates as the area gets bigger. This region of supersonic acceleration is terminated by a normal shock wave. The shock wave produces a near-instantaneous deceleration of the flow to subsonic speed. This subsonic flow then decelerates through the remainder of the diverging section and exhausts as a subsonic jet. In this regime if the back pressure is lowered or raised the length of supersonic flow in the diverging section before the shock wave increases or decreases.
REVIEW: OPERATION OF C-D NOZZLES • If pb is lowered enough the supersonic region may be extended all the way down the nozzle until the shock is sitting at the nozzle exit, figure (d). Because of the very long region of acceleration (the entire nozzle length) the flow speed just before the shock will be very large. However, after the shock the flow in the jet will still be subsonic. • Lowering the back pressure further causes the shock to bend out into the jet, figure (e), and a complex pattern of shocks and reflections is set up in the jet which will now involve a mixture of subsonic and supersonic flow, or (if the back pressure is low enough) just supersonic flow. Because the shock is no longer perpendicular to the flow near the nozzle walls, it deflects it inward as it leaves the exit producing an initially contracting jet. We refer to this as over-expanded flow because in this case the pressure at the nozzle exit is lower than that in the ambient (the back pressure)- i.e. the flow has been expanded by the nozzle too much. • A further lowering of the back pressure changes and weakens the wave pattern in the jet. Eventually, the back pressure will be lowered enough so that it is now equal to the pressure at the nozzle exit. In this case, the waves in the jet disappear altogether, figure (f), and the jet will be uniformly supersonic. This situation, since it is often desirable, is referred to as the ‘design condition’, Pe=Pa.
REVIEW: OPERATION OF C-D NOZZLES • Finally, if the back pressure is lowered even further we will create a new imbalance between the exit and back pressures (exit pressure greater than back pressure), figure (g). In this situation, called under-expanded, expansion waves that produce gradual turning and acceleration in the jet form at the nozzle exit, initially turning the flow at the jet edges outward in a plume and setting up a different type of complex wave pattern. • Summary Points to Remember: • When flow accelerates (sub or supersonically) pressure always drops • Pressure rises instantaneously across a shock • Pressure falls across an expansion wave • Pressure throughout jet is always same as ambient (i.e. the back pressure) unless jet is supersonic and there are shocks or expansion waves in jet to produce pressure differences
KEY EQUATIONS FOR NOZZLE DESIGN Nozzle area ratio as a function of engine parameters Once nozzle area is set, operating point of engine depends only on tt Compare with Section 6.7 H&P A7 is the throat area, how do we find the exit area of the nozzle? Found from compressible channel flow relations, recall that M7=1 Set by jet stagnation pressure and ambient Compare with Equation 3.15
ROCKET NOZZLE BEHAVIOR • Low Altitude • OVER-EXPANDED • Pe < Pa • Do not expand beyond Pe=0.4 Pa • Intermediate Altitude • IDEALLY-EXPANDED • Pe = Pa • High Altitude • UNDER-EXPANDED • Pe > Pa
SUMMARY • Sea Level: Over-Expanded • Rockets operate at this condition at take-off • Intermediate Altitude: Ideally-Expanded • Typically occurs at only 1 point in rocket flight • High Altitude: Under-Expanded
EXAMPLE: SPACE SHUTTLE MAIN ENGINE, e~77 • Static pressure at exit of Space Shuttle Main Engine nozzle is considerably less than ambient pressure at sea level • Mismatch in pressure gives rise to Mach “disc” in nozzle exhaust • Extremely strong shock wave that creates a region of subsonic flow and produces a characteristic white luminescent glow • Is this nozzle Over- or Under-Expanded? • If this nozzle were to operate at this altitude only, should it be larger or smaller?
VARIABLE GEOMETRY AND THRUST VECTORING PW119 MIG-29 OVT F up to 15° in any direction http://www.pr.afrl.af.mil/divisions/prt/ihptet/brochure/Exhaust.htm
NEW TECHNOLOGY (?): Messerschmitt Me 262 World's first operational jet-powered fighter
SUMMARY: INLETS AND NOZZLES Expression for thrust that shows dependence on atmospheric pressure and cross-sectional area at compressor or fan entrance Engine Characteristics Nozzle area ratio as a function of engine parameters Once nozzle area is set, operating point of engine depends only on tt
COMMENTS ON NOISE REDUCTION USING NOZZLES • Environmental issues are likely to impose fundamental limitation on air transportation growth in the 21st century • 2 major contributors: NOISE and EMISSIONS • Noise • Primarily exhaust jet and fan (whirl) noise • Noise impact of subsonic aircraft is constrains air transportation system through curfews, noise budgets and slot restrictions • Some solutions • Exhaust mixers • Liners that absorb sound • Shaping of stators and fan blade components for low noise
WHY MIX CORE AND FAN FLOWS? Large velocity gradient Large shear tends to be noisy Acoustic noise level scales as U6jet-U8jet
INFLUENCE OF BYPASS RATIO:Relative Perceived Noise Level vs. Bypass Ratio • Increasing bypass ratio increases the fan noise moderately • Increasing bypass ratio decreases the jet noise dramatically
CORE-BYPASS MIXING Mixing of two streams of different Tt increases hprop How does level of static p at which mixing occurs change mixed-out Pt
NOISE CERTIFICATION REQUIREMENTSJET AND TRANSPORT AIRPLANES AT TAKEOFF
RADIAL IN-FLOW NOZZLES • Often referred to as spike nozzles • Named for prominent spike centerbody • May be thought of as a bell turned inside out • Nozzle is only one of many possible spike configurations • (a) traditional curved spike with completely external supersonic expansion • (b) similar shape in which part of the expansion occurs internally • (c) design similar to E-D nozzle in which all expansion occurs internally
SPIKE NOZZLES • Each of spike nozzles features a curved, pointed spike • Most ideal shape • Spike shape allows exhaust gases to expand through isentropic process • Nozzle efficiency is maximized and no energy is lost because of turbulent mixing • Isentropic spike may be most efficient but tends to be prohibitively long and heavy • Replace curve shape by shorter and easier to construct cone ~1% performance loss
AEROSPIKE NOZZLES • Further subclass of radial in-flow family of spike nozzles is known as aerospike • Go even further by removing pointed spike altogether and replace with a flat base • This configuration is known as a truncated spike
TRUNCATED AEROSPIKE NOZZLES • Disadvantage of "flat" plug is turbulent wake forms aft of base at high altitudes resulting in high base drag and reduced efficiency • Alleviated by introducing a "base bleed," or secondary subsonic flow • Circulation of this secondary flow and its interaction with the engine exhaust creates an "aerodynamic spike" that behaves much like the ideal, isentropic spike • Secondary flow re-circulates upward pushing on base to produce additional thrust • It is this artificial aerodynamic spike for which the aerospike nozzle is named
LINEAR AEROSPIKE NOZZLE • Still another variation of aerospike nozzle is linear (instead of annular) • Linear Aerospike pioneered by Rocketdyne (now division of Boeing) in 1970’s • Places combustion chambers in a line along two sides of nozzle • Approach results in more versatile design • Use of lower-cost modular combustors • Modules can be combined in varying configurations depending on application.
