Most subsonic aircraft have their engines placed in nacelles; thus, in this section we do not deal with the inlet alone but include the nacelle at subsonic Mach numbers. The cross section of a typical subsonic inlet and its geometric parameters are shown in Fig. 10.1. The inlet area A1 is based on the flow cross section at the inlet highlight. Because the subsonic inlet can draw in airflow whose free stream area A0 is larger than the inlet area A1, variable inlet geometry is not
required (except sometimes blow-in doors or auxiliary inlets are used to reduce installation drag during takeoff). The material in this section on subsonic inlets is based on a fixed-geometry inlet.
The operating conditions of an inlet depend on the flight velocity and mass flow demanded by the engine. Figure 10.2 shows the streamline pattern for three typical subsonic conditions. Figure 10.2a shows acceleration of the fluid external to the inlet that will occur when the inlet operates at a velocity lower than the design value or at a mass flow higher than the design value.
Figure 10.2c shows deceleration of the fluid external to the inlet that will occur at a velocity higher than design or a mass flow lower than design.
A list of the major design variables for the inlet and nacelle includes the following:
1) Inlet total pressure ratio and drag at cruise
2) Engine location on wing or fuselage (avoidance of foreign-object damage, inlet flow upwash and downwash, exhaust gas reingestion, ground clearance)
3) Aircraft attitude envelope (angle of attack, yaw angle, cross-wind takeoff)
4) Inlet total pressure ratio and distortion levels required for engine operation
5) Engine-out wind milling airflow and drag (nacelle and engine)
6) Integration of diffuser and fan flow path contour
7) Integration of external nacelle contour with thrust reverser and accessories
8) Flow field interaction between nacelle, pylon, and wing
9) Noise suppression requirements
A diffuser is a component of a fluid flow system designed to reduce the flow velocity and thereby increase the fluid pressure. All turbo machines and many other flow systems incorporate a diffuser (ex. closed circuit wind tunnels, the duct between the compressor and burner of a gas turbine engine, the duct at exit from a gas turbine connected to the jet pipe, the duct following the impeller of a centrifugal compressor, etc.). Turbo machinery flows are, in general, subsonic (M < 1) and the diffuser can be represented as a channel diverging in the direction of flow (see Figure 2.13).
The basic diffuser is a geometrically simple device with a rather long history of investigation by many researchers. The long timespan of the research is an indicator that the fluid mechanical processes within it are complex, the research rather more difficult than might be anticipated, and some aspects of the flow processes are still not fully understood.
There is now a vast literature about the flow in diffusers and their performance. Only a few of the more prominent investigations are referenced here. A noteworthy and recommended reference, however, which reviews many diverse and recondite aspects of diffuser design and flow phenomena, is that of Kline and Johnson (1986).
The primary fluid mechanical problem of the diffusion process is caused by the tendency of the boundary layers to separate from the diffuser walls if the rate of diffusion is too rapid. The result of too rapid diffusion is always large losses in stagnation pressure. On the other hand, if the rate of diffusion is too low, the fluid is exposed to an excessive length of wall and fluid friction losses become predominant. Clearly, there must be an optimum rate of diffusion between these two extremes for which the losses are minimized. Test results from many sources indicate that an included angle of about 2q = 7deg gives the optimum recovery for both
two-dimensional and conical diffusers angle of about 2q = 7deg gives the optimum recovery for both two-dimensional and conical diffusers.
Diffuser performance parameters
The diffusion process can be represented on a Mollier diagram, Figure 2.12b, by the change of state from point 1 to point 2, and the corresponding changes in pressure and velocity from p1 and c1 to p2 and c2. The actual performance of a diffuser can be expressed in several different ways:
(i) as the ratio of the actual enthalpy change to the isentropic enthalpy change;
(ii) as the ratio of an actual pressure rise coefficient to an ideal pressure rise coefficient.
The supersonic inlet is required to provide the proper quantity and uniformity of air to the engine over a wider range of flight conditions than the subsonic inlet is. In addition, the nature of supersonic flow makes this inlet more difficult to design and integrate into the airframe. In supersonic flight, the flow is decelerated by shock waves that can produce a total pressure loss much greater than, and in addition to, the boundary-layer losses. Indeed, at M0 = 3, a simple pitot-type inlet preceded by a normal shock wave would have a total pressure recovery hr of 0.32 due to the normal shock alone!
An engine overall compression ratio is the product of the engine’s ram, diffuser, and compressor pressure ratios:
thrust per unit mass flow are very sensitive to the diffuser pressure ratio Pd. For supersonic cruise flight, therefore, the design of the inlet becomes of paramount importance. For this reason, we shall now examine the basic principles and operating characteristics of supersonic aircraft inlets (or diffusers). The study of supersonic inlets is not new, and many excellent books and reports have been written for the benefit of students and practicing professionals (see Refs. 54-59 and 61-67). Special attention should be paid to Ref. 55, a textbook that covers the aerodynamics of inlets.