16.50 Lecture 22: Inlets or Diffusers

16.50 Lecture 22 Subjects: Inlets or Diffusers While the Gas Generator, composed of the compressor, combustor and turbine, is the heart of any gas turbine engine, the overall performance of the propulsion system is strongly influenced by the inlet and the nozzle. This is especially true for high Mo flight, when a major portion of the overall temperature and pressure rise of the cycle are in the inlet, and a correspondingly large part of the expansion in the nozzle. So it is important to understand how these components function and how they limit the performance of the propulsion system. Inlets or Diffusers These two titles are used interchangeably for the component that captures the oncoming propulsive streamtube and conditions it for entrance to the engine. The function of the inlet is to adjust the flow from the ambient flight condition, to that required for entry into the fan or compressor of the engine. It must do this over the full flight speed range, from static (takeoff) to the highest Mo the vehicle can attain. Comparing the simple diagrams of subsonic and supersonic inlets, we can appreciate that the subsonic inlet has the simpler task: From Kerrebrock, Jack L. Aircraft Engines and Gas Turbines. 2nd edition. MIT Press, 1992. © Massachusetts Institute of Technology. Used with permission. For any inlet the requirements are two: a) To bring the inlet flow to the engine with the highest possible stagnation pressure. This is measured by the Inlet Pressure Recovery, πd . b) To provide the required engine mass flow. As we shall see the mass flow can be limited by choking of the inlet. 1 Subsonic Inlets: The subsonic inlet must satisfy two basic requirements: a) Diffusion of the free-stream flow to the compressor inlet condition at cruise. b) Acceleration of static air to the compressor inlet condition at takeoff. There is a compromise to be made because a relatively thin "lip" aligned with the entering flow is best for the cruise condition. This is to avoid accelerating the flow, already at a Mach number of the order of 0.8, to supersonic speeds that will lead to shock losses. But a more rounded lip will better avoid separation for the takeoff condition because the air must be captured from a wide range of angles: The minimum area, A1 is set to just avoid choking (M = 1) in the inlet when M2 has its largest value (this is set by the maximum blade Mach number and the blade shape). A well designed subsonic inlet will produce a stagnation pressure recovery πd in the order of 0.97 at its design condition. A1 From Kerrebrock, Jack Aircraft L. Engines and Gas Turbines . 2nd edition. MIT Press, 1992. © Massachusetts Institute of Technology. Used with permission. Supersonic Inlets: At supersonic flight speeds the pressure and temperature rise in the inlet can be quite large. For the best thermodynamic efficiency it is important that this compression be as nearly reversible (isentropic) as possible. At a flight Mach number of 3 the ideal pressure ratio is ! p 2 ! "1 δο ≡ to = (1+ 2 Mo ) ! "1 =36.7 po while the temperature rise is T ! "1 2 θo = to = 1+ 2 Mo =2.8 To For the turbojet cycle the compression is partly in the inlet and partly in the compressor: 2 Note: T3-T2=T4-T5 From Kerrebrock, Jack L. Aircraft Engines and Gas Turbines. 2nd edition. MIT Press, 1992. © Massachusetts Institute of Technology. Used with permission. If the diffusion from point 0 to point 2 is not reversible, the entropy increases, and this results in a lower value of Pt2. This has two effects: a) The expansion ratio of the nozzle is decreased, so the jet velocity and thrust are lower. b) The lower pressure at point 2 limits the mass flow through the compressor to a lower value than for isentropic diffusion. So there is a double penalty for losses in the inlet. Unfortunately some losses are inevitable, the more so the larger Mo. To see why we will discuss the flow in supersonic diffusers, beginning with the simplest. 3 Normal-Shock diffuser: All existing compressors and fans require subsonic flow at their inlet with 0.5 < M2 < 0.8 at high power conditions. So the inlet must reduce the flow Mach number from Mo > 1 to M2 < 1. The simplest way to do this is with a Normal Shock. Here M1 < 1 is entirely determined by Mo, according to the normal shock relation M1 2 = 1+ ! 2"1 Mo 2 !Mo 2 " !2"1 The stagnation and static pressures are also determined by Mo: 1 2! Mo 2 ! " 1 ! "1 ( ! " 1)Mo 2 + 2 ! !" 1 pto =( " ) ( 2 ) pt1 ! +1 ! +1 (! +1)Mo p1 2! 2 = 1+ (M 0 "1) p0 ! +1 From Kerrebrock, Jack L. Aircraft Engines and Gas Turbines. 2nd edition. MIT Press, 1992. © Massachusetts Institute of Technology. Used with permission. For low supersonic speeds, such diffusers are adequate because the stagnation pressure loss is small, but at Mo = 2, pt2/pto ≈ .71, a serious penalty, and at Mo = 3 p2t/pto ≈ .32. For example the F-16 fighter has a simple normal shock diffuser, while the F-15 has an oblique shock diffuser such as will be discussed next. 4 Oblique - Shock diffusers: The losses can be greatly reduced by decelerating the flow through one or more oblique shocks, the deflection and the pressure rise of each being small enough to be in the range where the stagnation pressure ratio is close to unity. It is very important to understand that an Oblique Shock is in fact just a normal shock standing at an angle to the flow. All of the change across the shock takes place because the Mach number normal to the shock is larger than unity. The velocity normal to the shock decreases with consequent increases in temperature and pressure, but the velocity parallel to the shock is unchanged: M1n is given in terms of Mon by the same relation given for M1 as a function of Mo. But Mon can be made close to 1. The condition for a weak or sound wave is just Mon = 1, or 1 Sinθ = M (weak compression) o By choosing the wedge angle (or deflection angle) δ we can set the shock angle. A series of weak oblique shocks, for each of which the Mn is near unity, hence all lying in the range of small pt loss, can yield an efficient diffuser. 5 Diffusers with internal contraction One might ask why we do not just use a convergent-divergent nozzle in reverse as an inlet: This would work at one design Mach number, the one for which the isentropic area ratio between the incoming supersonic flow and the sonic throat is exactly the asbuilt area ratio A1/Athroat . But during the acceleration to this Mach number the fully supersonic flow cannot be established in the inlet without varying the geometry. To see this, imagine the inlet flying at Mo , lower than the design Mach number. The flow will look as depicted in the top right diagram below. This is because at the lower M0 the flow area that would decelerate isentropically to sonic at the throat is smaller than the built A1: A/A* A1/A* A0/A* 0 1 M0 1.5 the loss due to the shock at the throat is still excessive. 7