INLET AND EXAUST NOZZLES Chap. 10 AIAA AIRCRAFT ENGINE DESIGN R01-07/11/2011

Transcription

1 MASTER OF SCIENCE IN AEROSPACE ENGINEERING PROPULSION AND COMBUSTION INLET AND EXAUST NOZZLES Chap. 10 AIAA AIRCRAFT ENGINE DESIGN R01-07/11/2011 LECTURE NOTES AVAILABLE ON Prof. Eng. Antonio Ficarella 1

2 CONCEPTS integration of the engine and airframe engine performance (offdesign) cycle analysis major impact on the performance inlet area vs. Mach number and altitude partial throttle conditions inlet total pressure ratio vs. Mach n. 2

3 INLETS overall pressure ratio πr freestream pressure recovery πd diffuser total pressure ratio πc compressor pressure ratio πi = πr πd 3

4 inlet: compressor with no moving parts adiabatic efficiency τr = τi 4

5 inlet polytrophic efficiency the inlet is superior to mechanical compressor up to Mach 3.5 5

6 PERFORMANCE OF AN INLET to bring the required air with min pressure loss and flow distortion controllable flow matching low installation drag good starting and stability low acoustic, radar and infrared signature min weight and cost life and reliability goals sub and supersonic inlets differ considerably efficient and stable supersonic diffusion over a wide range of Mach n. is very difficult to achieve variable geometry 6

7 DESIGN TOOLS SUBSONIC INLETS because the subsonic inlet can draw in airflow whose freestream area (A0) is larger than the inlet area (A1), variable inlet geometry is not required except that blow-in doors during takeoff 7

8 inlet total pressure ratio πd was assumed to be constant and equal to πdmax the total pressure ratio due to friction 8

9 the diameter at the throat Dth is sized such that the Mach n. does not exceed 0.8 freestream area of chocked flow at the sea level A*0 9

10 the AOA can vary (takeoff and landing) angle of attack (AOA): velocity of the air (V) to the wing chord line (WCL) the airflow angle with respect to the engine centerline will change inclining the face of the nacelle 10

11 effect: flow separation Dhl/Dth =

12 INLET FLOW DISTORTION 12

13 INET DRAG AND ENGINEOUT DRAG drag-divergence Mach n. Mdd at which the drag increase dramatically 13

14 14

15 NACELLE AND INTERFERENCE DRAG 15

16 DIFFUSER maintaining the flow attached to the inside walls in the presence of the adverse pressure gradient, boundary layers tend to separate VORTEX GENERATORS 16

17 17

18 ZERO FLIGHT SPEED sharp curvature of the streamlines flow separation 18

19 DESIGN TOOLS SUPERSONIC INLETS the nature of supersonic flow makes this inlet more difficult to design GASTAB can calculate the change in properties across the shocks 19

20 INLET TYPES INTERNAL COMPRESSION EXTERNAL COMPRESSION MIXED COMPRESSION 20

21 INTERNAL COMPRESSION a series of internal oblique shocks waves followed by a terminal shock downstream of the throat requires variable throat area to swallow the normal shock during starting and fast bypass doors unstarted inlet flow pattern distrupted Aths to start the inlet / Athr at normal operation Mth=1.2 many experts do not include it as a useful type of inlet 21

22 EXTERNAL COMPRESSION INLET a series of oblique shocks followed by a normal shock 22

23 pitot inlet acceptable only for Mach < 1.6 the drag at the inlet is associated mainly with the loss of momentum of the excess air captured 23

24 WITH OBLIQUE SHOCKS the external compression ramp turns the flow diffuser duct must turn back the flow which may add weight, length, friction ACCEPTABLE MACH <

25 MIXED COMPRESSION INLET at flight Mach n. above 2.5 two-dimensional (rectangular) and axisymmetric (circular) 25

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27 TOTAL PRESSURE RECOVERY ηr estimated by 27

28 for a general inlet compression inlet 28

29 total pressure ratio across a normal shock Mx upstream My downstream Mach n. for oblique shocks Mx replaced by M1sinβ (M1 upstream) and My replaced by M2sin(β- θ) 29

30 two limiting values θ* leading to M2=1 θmax for which an oblique shock solution exist 30

31 MASS FLOW CHARACTERISTICS inlet mass flow ratio the ratio of the actual mass flow rate to the mass flow rate captured by the undisturbed geometrical opening difference: air spilled around the inlet engine mass flow ratio difference boundary layer bleed 31

32 inlet matched to the engine critical operation terminal shock just inside the lip min air fraction spilled at the inlet UNMATCHED supercritical operation the shock is sucked down into the diffuser lower inlet total pressure recovery 32

33 BUZZ STABILITY OF TERMINAL SHOCK low frequency, high amplitude pressure oscillation Mach throat 1.2 normal shock downstream where Mach 1.3 UNSTARTING expelling the normal shock when engine needs less air excess air must bypass the engine 33

34 BOUNDARY LAYER SEPARATION there is an adverse pressure gradient 34

35 INLET DESIGN AND SIZING the capture and throat areas must be large enough not to choke the airflow required by the engine for supersonic flight conditions, the inlet s capture area A1 is sized to capture the required air flow since the airflow varies with both flight Mach and engine throttle, variable design geometry is needed 35

36 INLET DRAG 36

37 REQUIRED INLET AIR FLOW required engine airflow m0 (m0spec) and the corresponding freestream area A0 (A0spec) are based on the calculation of ηrspec since ηr may be different from ηrspec 37

38 the engines operates as a constant mass flow device mc2=constant 4% margin for boundary layer bleed 38

39 INLET MASS FLOW 39

40 INLET SIZE ref = reference cycle require inlet size at a flight condition 40

41 INLET AIRFRAME INTERFERENCE EFFECTS 41

42 SUBSONIC DIFFUSER 42

43 EXISTING INLET DESIGN FIXED, DOUBLE RAMP 43

44 VARIABLE, TRIPLE RAMP 44

45 FLOW SEPARATION FROM SHARP LIP INLETS blow-in doors to reduce additive drags (takeoff) 45

46 EXAUST NOZZLES to increase the velocity of exhaust gases to collect and straighten the flow the pressure ratio controls the expansion process max thrust when the exit pressure Pe equals the ambient pressure P0 NOZZLE ADIABATIC EFFICIENCY es = ideal exit conditions if inlet kinetic energy is small nozzle polytrophic efficiency 46

47 47

48 NOZZLE TYPES CONVERGENT NOZZLE when pressure ratio is low (<4) CONVERGENT-DIVERGENT (C-D) NOZZLE variable geometry 48

49 EJECTOR-NOZZLE CONFIGURATION 49

50 NOZZLE FUNCTIONS back-pressure control for the engine thrust vectoring ENGINE BACK PRESSURE CONTROL at reduced engine corrected mass flow rates (reduced throttle), the operating line of a compressor move closer to stall line INCREASE THE NOZZLE AREA to reduce the engine back-pressure and increase the corrected mass flow through the compressor 50

51 afterburning engines large change in nozzle throat area opening the nozzle to max area reduces the turbine back pressure the necessary power for starting may be produced at lower turbine temperature and the compressor may be started at a lower speed reduces the size of the starter 51

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53 THRUST REVERSING AND VECTORING 53

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55 DESIGN TOOLS EXHAUST NOZZLES TOTAL PRESSURE LOSS two-dimensional or axisymmetric nozzle the total pressure loss of a rectangular nozzle with an aspect ratio AR (W/H) of 2 will be 1.2 times that of a circular nozzle of the same area 55

56 NOZZLE COEFFICIENTS gross thrust coefficient discharge or flow coefficient if A8e is the actual area required to pass the flow of engine cycle calculation: 56

57 57

58 velocity coefficient 58

59 angular coefficient CA thrust loss due to the non-axial exit 59

60 ONE-DIMENSIONAL FLOW for adiabatic flow CA=1 60

61 separation: effective exit pressure just preceding the shock wave (0.37P0) there may be insufficient thrust increase for a C-D nozzle on a subsonic cruise to pay for the additional drag and weight of such a nozzle 61

62 62

63 GENERAL THRUST PERFORMANCE gross thrust coefficient 63

64 neglecting leakage and cooling 64

65 OFF-DESIGN 65

66 EXAMPLES 66