Measurement of the pressure at various points on the surface of the airfoil will reveal a pressure distribution as shown in Figure 7-5c. The vectorial sum of these pressures will produce some resultant force acting on the blade. This resultant force can be resolved into a lift component Lat right angles to the undisturbed air stream， and a drag component D， moving the airfoil in the direction of flow motion. This resultant force is assumed to act through a definite point located in the airfoil so that the behavior will be the same as if all the individual components were acting simultaneously.
By experimentation， it is possible to measure the lift and drag forces for allvalues of airflowvelocity， angles of incidence， various airfoil shapes.Thus， for any one airfoil the acting forces can be represented as shown in Figure7-6a. Using such observed values， it is possible to define relations between the forces
V2 D二 CD Aρ 2 (7-2)
V2 L二 CL Aρ 2 (7-3)
where:
L二 lift force
D二 drag force
CL二 lift coefficient
CD二 drag coefficient
A二 surface area
ρ二 fluid density
V二 fluid velocity
Two coefficients have been defined， CL and CD， relating velocity，density，area， and lift or drag forces. These coefficients can be calculated from wind-tunnel tests and plotted as shown in Figure 7-6b versus the angle of attack for any desired section. These curves can then be employed in all future predictions involving this particular foil shape.
Examination of Figure 7-6b reveals that there is an angle of attack thatproduces the highest lift force and lift coefficient. If this angle isexceeded， the airfoil ""stalls"" and the drag force increases rapidly. As this maximumangle is approached， a great percentage of the energy available is lost inovercoming friction， and a reduction in efficiency occurs. Thus， there is apoint， usually before the maximum lift coefficient is reached， at which the most economical operation occurs as measured by effective lift for a given energy supply.
.aminar-Flow Airfoils
.ustbefore and during World WarII， much attention was given to laminar-flow airfoils. These airfoils are designed so that the lowest pressure on the surface occurs as far back as possible. The reason for this design is that the stability of the laminar boundary layer increases when the external flow is accelerated (in the flow with a pressuredrop)， and the stability decreases when the flow is directed against increasing pressure. A consider-able reduction in skin friction is obtained by extending the laminar region inthisway， provided that the surface is sufficiently smooth.
A disadvantage of this type of airfoil is that the transition from laminar to turbulent flow moves forward suddenly at small angles of attack. Thissudden movement results in a narrow low-drag bucket， which means that the drag at moderate-to-large attack angles is much greater than an ordinary airfoil for the same attack angle as seen in Figure 7-7. This phenomenon canbe attributed to the minimum pressure point moving forward;therefore， the point of transition between laminar and turbulent flow is also advanced toward the nose as shown in Figure 7-. The more an airfoil is surroundedby turbulent airflow， the greater its skin friction.
Cascade Test
The data on blades in an axial-flow compressor are from various types ofcascades， since theoretical solutions are very complex， and their accuracy is in question because of the many assumptions required to solve the equations. The most thorough and systematic cascade testing has been conducted by NACA staff at the Lewis Research Center. The bulk of the cascade testing was carried out at low mach numbers and at low turbulence levels.
The NACA 65 blade profiles were tested in a systematic manner byHerrig，Emery， and Erwin. The cascade tests were carried out in a cascade wind tunnel with boundary-layer suction at the end walls. Tip effects were studied in a specially designed water cascade tunnel with relative motion between wall and blades.
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