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This review analyzes key components of aircraft propulsion engines, including diffusers, nozzles, compressors, and turbines for Aerospace Engineering students. It covers concepts like ideal and non-ideal diffusers, supersonic flows, shock formations, pressure rise limits, and oblique shocks. The analysis addresses operational modes, design conditions, and off-design scenarios to enhance understanding of engine performance. Constructed as a comprehensive reference, this review sheds light on critical aspects of propulsion systems for academic exploration.
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Review ofComponents Analysis Aerospace Engineering, International School of Engineering (ISE) Academic year : 2012-2013 (August – December, 2012) Jeerasak Pitakarnnop , Ph.D. Jeerasak.p@chula.ac.th jeerasak@nimt.or.th Aircraft Propulsion
Component Analysis • Diffuser • Free Stream to Diffuser Inlet • Diffuser Inlet to Outlet • Nozzle • Fixed Divergent Nozzle • Diverging Converging Nozzle • Axial Flow Compressor • Axial Flow Turbine Aircraft Propulsion
Engine without Inlet Cone • Free Stream to Diffuser Inlet • Subsonic Flow • Supersonic Flow with Shock • Diffuser Inlet to Outlet • Ideal Diffuser – Isentropic Flow • Non Ideal Diffuser – Fanno Line Flow Aircraft Propulsion
Free Stream to Diffuser Inlet πo represents loss from free stream to the inlet. Supersonic Flow Shock • πo < 1 Subsonic Flow • πo ≈ 1 (= 1: ideal isentropic flow) Aircraft Propulsion
Supersonic Flow with Normal Shocks • Shocks usually occur exterior to, or near, the inlet plane of the diffuser when an aircraft flies supersonically. • The strongest shocks is the normal shocks. Aircraft Propulsion
Ex 1: Normal Shocks A standing normal shock occurs on an aircraft flying at Mach 1.50. The internal recovery factor of the diffuser is 0.98, and the specific heat ratio is 1.40. Find the total recovery factor of the diffuser. Aircraft Propulsion
Ideal Diffuser Isentropic & Adiabatic Flow • Constant Total Pressure pta = pt1 = pt2 • Constant Total Temperature Tta= Tt1= Tt2 (hta= ht1= ht2) Aircraft Propulsion
Isentropic Flow Mach Number and Local Speed of Sound Stagnation Relations Area Ratio Aircraft Propulsion
Limit on Pressure Rise Separation is one of the limits of the diffuser operation. Aligned Inlet Flow: for flow without separation. Mis-Aligned Inlet Flow: Upper limit on the pressure coefficient will be reduced appreciably to perhaps 0.1 to 0.2. Aircraft Propulsion
Ex 2: Separation Limit Design an ideal diffuser to attain the maximum pressure rise if the incoming Mach no. is 0.8. That is find the diffuser area ratio, pressure ratio and the resulting exit Mach number. Assuming isentropic flow and γ = 1.4. Aircraft Propulsion
Non Ideal Diffuser Low Speed/Flow accelerates/Pressure decreases To quantify loss from the free stream to the diffuser exit, we introduce: • Total Pressure Recovery Factor: High Speed/Flow decelerates/ Pressure increases Nearly Adiabatic Flow, assume: • Constant Total Enthalpy hta= ht1= ht2 • Constant Total Temperature Tta = Tt1 = Tt2 where • πr is the diffuser pressure recovery factor, and • πo represents loss from free stream to the inlet. Aircraft Propulsion
Friction Flow Viscous flows are the primary means by which total pressure losses occur!! Fanno Line Flow: flow with friction but no heat transfer Fanno Line Flow could be used when: • Exit-to-inlet area ratio is near unity, • The flow does not separate. Aircraft Propulsion
Fanno Line Flow Adiabatic Flow of a Calorically Perfect Gas in a Constant-Area Duct with Friction Aircraft Propulsion
Engine with Inlet Cone • Oblique Shock • Oblique Planar Shock • Oblique Conical Shock • Mode of Operation • Design Condition • Off Design Condition Aircraft Propulsion
Oblique Planar Shocks • 2D planar shock is simpler than conical shock. • Occur when an inlet is attached to the fuselage of the aircraft, the inlet is more or less rectangular, resulting in planar shock. • Flow behind the planar shock is uniformly parallel to the wedge. Aircraft Propulsion
Oblique Planar Shocks δ = deflection angle σ = shock angle Aircraft Propulsion
Oblique Conical Shocks • Found in many aircraft applications. • A conical ramp is used to generate an oblique shock, which decelerate flow to a less supersonic conditions. • A normal shock further decelerates the flow to a subsonic condition for the internal flow in the diffuser. Spike on BlackBird Aircraft Propulsion
Oblique Conical Shocks Aircraft Propulsion
Oblique Conical Shocks Aircraft Propulsion
Oblique Conical Shocks Aircraft Propulsion
Modes of Operation Design Condition: the oblique shock intersects the diffuser cowl All the air that cross oblique shock enters the engine Flow rate decreases Pressure in the diffuser decreases Mach no. in the diffuser decreases Shock is pushed out!! Shock is stronger larger total pressure loss Some of the air will be spilled high pressure additive drag Shock is used to compress air outside shock wasting power Flow rate increases Pressure in the diffuser drops Shock moves into the diffuser Acting like a supersonic nozzle Shock occurs in diverging section with high Mach no. More total pressure is lost. Aircraft Propulsion
Mass Flow or Area Ratio Reference Parameter True ingested mass Mass flow enters the engine Mass flow ratio Aircraft Propulsion
Design Operation Aircraft Propulsion
Off-Design Operation When the diffuser operated at off-design conditions, the area should be varied so that it operates efficiently. In the case of a single planar oblique shock: Inlet area could be determined from: Aircraft Propulsion
Ex 3: Supersonic Diffuser A diffuser with a spike is used on a supersonic aircraft. The freestream Mach number is 2.2, and the cone half-angle is 24°. The standing oblique shock is attached to the spike and cowl, and a converging inlet section with a throat of area Am is used to decelerate the flow through internal compression. Assume γ = 1.4 and πr = 0.98. • Estimate πd on the assumption the inlet starts. Also, find the required Am/A1 • Find πd on the assumption the inlet doesn’t start and has a standing normal shock located in front of the spike. Aircraft Propulsion
Nozzle Fixed Diverging Nozzle Converging-Diverging Nozzle Aircraft Propulsion
Primary Nozzle In real analysis: • Pexit may not match Pa due to incorrect nozzle area proportion. • Frictional losses are include but adiabatic process still be assumed. Nozzle Efficiency Constant cp Specific heat Adiabatic Exit Velocity Aircraft Propulsion
Primary Nozzle Adiabatic Process Flow: For the ideal case, isentropic process Thus, Adiabatic Aircraft Propulsion
Primary Nozzle Choke condition: If p* > pa, the nozzle is choke If p*= p8, M8 = 1 If p* < pa , M8 < 1 and p8 = pa Then, Aircraft Propulsion
Converging Nozzle Exhaust of converging nozzle with matching exhaust and ambient pressures Exhaust of under expanded converging nozzle Aircraft Propulsion
Converging-Diverging Nozzle Aircraft Propulsion
1st – 3rd Condition of CD nozzle • Case 1: pexhaust = pambient and Subsonic Flow Through out the nozzle. • Case 2: pexhaust = pambient and Subsonic Flow Through out the nozzle but Mthroat =1. • Case 3: pexhaust = pambient ,Subsonic Flow in the converging section and Supersonic Flow in the diverging section. • MAXIMUM THRUST • Design Condition for the ideal case Aircraft Propulsion
4th Condition of CD Nozzle • pambientis slightly above the designed pexhaust • Result in a complex 2D flow pattern outside the nozzle • Considered as “Overexpanded Case” • The flow suddenly is compressed and decelerates outside the nozzle. • A series of compression waves and expand waves are generated. • Can be calculated basing on 2D compressible flow Aircraft Propulsion
5th Condition of CD Nozzle • pambientis below the designed pexhaust • Result in a complex 2D flow pattern outside the nozzle • Considered as “UnderexpandedCase” or “Super Critical Case” • The flow continues to expand and accelerates outside the nozzle. • A series of compression waves and expanded waves are generated resulting in a series of shock diamonds. Aircraft Propulsion
6th Condition of CD Nozzle • pambientis significantly above the designed pexhaustBut below the 2nd case • Result in a single normal shock or a series of oblique and normal shocks called λ • Also “Overexpanded Case” • Result in a subsonic exit Mach no.: LOW THRUST Totally undesirable Aircraft Propulsion
7th Condition of CD Nozzle • pambientis significantly above the designed pexhaustLimit condition of the 6th case • Exit pressure causes a normal shock exactly at the exit plane • Case 4 falls between case 7 and 3. Aircraft Propulsion
Ex4: Converging-Diverging Nozzle A converging-diverging nozzle with an exit area of 0.2258 m2 and a minimum area of 0.1774 m2 has an upstream total pressure of 137.895 kPa. The nozzle efficiency is 0.965 and the specific heat ratio is 1.35. • At what atmospheric pressure will the nozzle flow be shockless? • At what atmospheric pressure will a normal shock stand in the exit plane? Aircraft Propulsion
Axial Flow Compressor Aircraft Propulsion
Velocity Polygon Aircraft Propulsion
Total Pressure Ratio The equations is derived for a single stage (rotor and stator) using 2D planar mean line c.v. approach. “Midway between hub and tip” • Power Input to the Shaft • Total Pressure Ratio of the Stage Control Volume definition for compressor stage Aircraft Propulsion
Percent Reaction A relation that approximates the relative loading of the rotor and stator based on the enthalpy rise: Aircraft Propulsion
Relationships of Velocity Polygons to Percent Reaction and Pressure Ratio Aircraft Propulsion
Limit on Stage Pressure Ratio • The rotor is moving, the relative velocity must be used: • For the stator, which is stationary the relative velocity must be used: 1 and 2 refer to the stage inlet and midstage properties. Aircraft Propulsion
Limit on Stage Pressure Ratio Stator Rotor Aircraft Propulsion
Axial Flow Turbine Aircraft Propulsion
Velocity Polygon Aircraft Propulsion
Velocity Polygon Aircraft Propulsion
Total Pressure Ratio The equations is derived for a single stage (rotor and stator) using 2D planar mean line c.v. approach. “Midway between hub and tip” The continuity, momentum and energy equations are used for the delivered shaft power: • Power Input to the Shaft • Total Pressure Ratio of the Stage Aircraft Propulsion
Percent Reaction A relation that approximates the relative loading of the rotor and stator based on the enthalpy rise: Aircraft Propulsion
Relationships of Velocity Polygons to Percent Reaction and Pressure Ratio Aircraft Propulsion