Mach angle

In the case of the finite projectile, the shock wave leaves the tip at an angle with the main flow which exceeds the Mach angle, on account of the conical nose which follows the tip. Appropriate corrections may be applied, however, and the shock-wave angle from such a sharp-nosed object remains an accurate means of measuring supersonic velocities. 

The incidence angle now must be corrected for the Mach number effect The effect of the Mach number on incidence angle is shown in Figure 7-26. The incidence angle is not affected until a Mach number of. 7 is reached. 

The incidence angle is now fully defined. Thus, when the inlet and outlet air angles and the inlet Mach number are known, the inlet blade angle can be computed in this manner. 

Carter s rule, which shows that the deviation angle is directly a function of the camber angle and is inversely proportional to the solidity 8 = mQ Xja) has been modified to take into account the effect of stagger, solidity, Mach number, and blade shape as shown in the following relationship ... 

Any effect of Mach number is experienced by rotor and stator equally and thus neither (or both) are limiting, and this Mach number will be lower than for other degrees of reaction under the conditions stated. If equal lift and drag are assumed in both rotor and stator, then optimum efficiency is obtained with R = 0.5 and VJu = 0.5. Although the latter is not always true, it does provide a useful criterion. Furthermore, the blade angles are similar in rotor and stator, which may be an advantage in the... 

In the expansion wave, the flow velocity is increased and the pressure, density, and temperature are decreased along the stream line through the expansion fan. Since Oj > 02, it follows that Mi flow through an expansion wave is continuous and is accompanied by an isentropic change known as a Prandtl-Meyer expansion wave. The relationship between the deflection angle and the Mach number is represented by the Prandtl-Meyer expansion equation.l - l... 

Fig. D-5 shows an external compression air-intake designed for optimized use at Mach number 2.0. Fig. D-6 shows a set of computed airflows of an external compression air-intake designed for use at Mach number 2.0 (a) critical flow, (b) sub-critical flow, and (c) supercritical flow. The pressures at the bottom wall and the upper wall along the duct flow are also shown. Two oblique shock waves formed at two ramps are seen at the tip of the upper surface of the duct at the critical flow shown in Fig. D-6 (a). The reflected oblique shock wave forms a normal shock wave at the bottom wall of the throat of the internal duct. The pressure becomes 0.65 MPa, which is the designed pressure. In the case of the subcritical flow shown in Fig. D-6 (b), the shock-wave angle is increased and the pressure downstream of the duct becomes 0.54 MPa. However, some of the airflow behind the obhque shock wave is spilled over towards the external airflow. Thus, the total airflow rate becomes 68% of the designed airflow rate. In the case of the supercritical flow shown in Fig. D-6 (c), the shock-wave angle is decreased and the pressure downstream of the duct becomes 0.15 MPa, at which the flow velocity is stiU supersonic.

In one vendor s design, a gas/wiWA. mixture enters each cell at a li IV< angle off vertical to imparl movement to one side of the macH r where the skimmer is located. 

The data collected during the test program consist of the fluid injection and wind tunnel parameters corresponding to the spark-shadow photographs. Water and air were used as the test fluids. Data were gathered for different sizes of injector nozzles at various levels of injection pressure and tunnel Mach number. Different injection angles also were examined. 

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