CONFIGURATION EFFECTS

TRANSONIC AND SUPERSONIC FLIGHT

Any object in subsonic flight which has some finite thickness or is producing lift will have local velocities on the surface which are greater than the free stream velocity. Hence, compressibility effects can be expected to occur at flight speeds less than the speed of sound. The transonic regime of flight pro­vides the opportunity for mixed subsonic and supersonic flow and accounts for the first | significant effects of compressibility.

Consider a conventional airfoil shape as shown in figure 3-9. If this airfoil is at a flight Mach number of 0.50 and a slight posi­tive angle of attack, the maximum local velocity on the surface will be greater than the flight speed but most likely less than sonic speed. Assume that an increase in flight Mach number to 0.72 would produce first evidence of local sonic flow. This condition of flight would be the highest flight speed possible without supersonic flow and would be termed the ‘ ‘critical Mach number. ’ ’ Thus, critical Mach number is the boundary between subsonic and transonic flight and is an im­portant point of reference for all compressi – | bility effects encountered in transonic flight. By definition, critical Mach number is the “free stream Mach number which produces first evidence of local sonic flow.” Therefore, shock waves, buffet, airflow separation, etc., take place above critical Mach number.

As critical Mach number is exceeded an area of supersonic airflow is created and a normal

Revised January 1965

M = .72

(CRITICAL MACH NUMBER)

M = .800 a = +2° CL= .442 SHOCK FORMATION IS APPARENT AT 25 TO 30 % CHORD POSITION

M = .875 a =4-2° CL= .450 SHOCK INDUCED SEPARATION ALONG AFT PORTION OF WING PLANFORM

shock wave forms as the boundary between the supersonic flow and the subsonic flow on the aft portion of the airfoil surface. The acceleration of the airflow from subsonic to supersonic is smooth and unaccompanied by shock waves if the surface is smooth and the transition gradual. However, the transition of airflow from supersonic to subsonic is always accompanied by a shock wave and, when there is no change in direction of the airflow, the wave form is a normal shock wave.

Recall that one of the principal effects of the normal shock wave is to produce a large increase in the static pressure of the airstream behind the wave. If the shock wave is strong, the boundary layer may not have sufficient kinetic energy to withstand the large, adverse pressure gradient and separation will occur. At speeds only slightly beyond critical Mach number the shock wave formed is not strong enough to cause spearation or any noticeable change in the aerodynamic force coefficients. However, an increase in speed above critical Mach number sufficient to form a strong shock wave can cause sepa­ration of the boundary layer and produce sudden changes in the aerodynamic force coefficients. Such a flow condition is shown in figure 3.9 by the flow pattern for M=0.77. Notice that a further increase in Mach number to 0.82 can enlarge the supersonic area on the upper surface and form an additional area of supersonic flow and normal shock wave on the lower surface.

As the flight speed approaches the speed of sound the areas of supersonic flow enlarge and the shock waves move nearer the trailing edge. The boundary layer may remain sepa­rated or may reattach depending much upon the airfoil shape and angle of attack. When the flight speed exceeds the speed of sound the ‘ ‘ bow’ ’ wave forms at the leading edge and this typical flow pattern is illustrated in figure 3 9 by the drawing for M=1.05. If the speed is increased to some higher supersonic value all oblique portions of the waves incline more greatly and the detached normal shock portion of the bow wave moves closer to the leading edge.

Of course, all components of the aircraft are affected by compressibility in a manner somewhat similar to that of basic airfoil. The tail, fuselage, nacelles, canopy, etc. and the effect of the interference between the various surfaces of the aircraft must be considered.

FORCE DIVERGENCE. The airflow sepa­ration induced by shock wave formation can create significant variations in the aerody­namic force coefficients. When the free stream speed is greater than critical Mach number some typical effects on an airfoil section are as follows :

(1) An increase in the section drag coeffi­cient for a given section lift coefficient.

(2) A decrease in section lift coefficient for a given section angle of attack.

0) A change in section pitching moment coefficient.

A reference point is usually taken by a plot of drag coefficient versus Mach number for a constant lift coefficient. Such a graph is shown in figure З. Ю. The Mach number which produces a sharp change in the drag coefficient is termed the “force divergence” Mach number and, for most airfoils, usually exceeds the critical Mach number at least 5 to 10 percent. This condition is also referred to as the “drag divergence” or “drag rise.” PHENOMENA OF TRANSONIC FLIGHT. Associated with the "drag rise” are buffet, trim and stability changes, and a decrease in control surface effectiveness. Conventional aileron, rudder, and elevator surfaces sub­jected to this high frequency buffet may “buzz,” and changes in hinge moments may produce undesirable control forces. Of course, if the buffet is quite severe and prolonged, structural damage may occur if this operation is in violation of operating limitations. When airflow separation occurs on the wing due to

shock wave formation, there will be a loss of lift and subsequent loss of downwash aft of the affected area. If the wings shock unevenly due to physical shape differences or sideslip, a rolling moment will be created in the direction of the initial loss of lift and con­tribute to control difficulty (‘‘wing drop”). If the shock induced separation occurs sym­metrically near the wing root, a decrease in downwash behind this area is a corollary of the loss of lift. A decrease in downwash on the horizontal tail will create a diving moment and the aircraft will ‘‘tuck under.” If these conditions occur on a swept wing planform, the wing center of pressure shift contributes to the trim change—root shock first moves the wing center of pressure aft and adds to the diving moment; shock formation at the wing tips first moves the center of pressure forward and the resulting climbing moment and tail

downwash change can contribute to ‘‘pitch up.”

Since most of the difficulties of transonic flight are associated with shock wave induced flow separation, any means of delaying or alleviating the shock induced separation will improve the aerodynamic characteristics. An aircraft configuration may utilize thin surfaces of low aspect ratio with sweepback to delay and reduce the magnitude of transonic force divergence. In addition, various methods of boundary layer control, high lift devices, vortex generators, etc., may be applied to improve transonic characteristics. For exam­ple, the application of vortex generators to a surface can produce higher local surface veloci­ties and increase the kinetic energy of the boundary layer. Thus, a more severe pressure gradient (stronger shock wave) will be neces­sary to produce airflow separation.

Once the configuration of a transonic air­craft is fixed, the pilot must respect the effect of angle of attack and altitude. The local flow I velocities on any upper surface increase with an increase in angle of attack. Hence, local sonic flow and subsequent shock wave formation can occur at lower free stream Mach numbers. A pilot must appreciate this reduction of force divergence Mach number with lift coefficient since maneuvers at high speed may produce compressibility effects which may not be en­countered in unaccelerated flight. The effect of altitude is important since the magnitude of any force or moment change due to com­pressibility will depend upon the dynamic pressure of the airstream. Compressibility effects encountered at high altitude and low dynamic pressure may be of little consequence in the operation of a transonic aircraft. How­ever, the same compressibility effects en­countered at low altitudes and high dynamic pressures will create greater trim changes, heavier buffet, etc., and perhaps transonic flight restrictions which are of principal inter­est only to low altitude.

PHENOMENA OF SUPERSONIC FLIGHT. While many of the particular effects of super­sonic flight will be presented in the detail of later discussion, many general effects may be anticipated. The airplane configuration must have aerodynamic shapes which will have low drag in compressible flow. Generally, this will require airfoil sections of low thickness ratio and sharp leading edges and body shapes of high fineness ratio to minimize the supersonic wave drag. Because of the aft movement of the aerodynamic center with supersonic flow, the increase in static longitudinal stability will demand effective, powerful control surfaces to achieve adequate controllability for super­sonic maneuvering.

As a corollary of supersonic flight the shock wave formation on the airplane may create special problems outside the immediate vicinity of the airplane surfaces. While the shock waves a great distance away from the airplane

can be quite weak, the pressure waves can be of sufficient magnitude to create an audible disturbance. Thus, “sonic booms" will be a simple consequence of supersonic flight.

The aircraft power pi ante for supersonic flight must be of relatively high thrust output. Also, in many cases it may be necessary to provide the air breathing powerplant with special inlet configurations which will slow the airflow to subsonic prior to reaching the compressor face or combustion chamber. Aero­dynamic heating of supersonic flight can pro­vide critical inlet temperatures for the gas turbine engine as well as critical structural temperatures.

The density variations in airflow may be shown by certain optical techniques. Schlieren photographs and shadowgraphs can define the various wave patterns and their effect on the airflow. The Schlieren photographs presented in figure 3-11 define the flow conditions on an aircraft in supersonic flight.