drag

This figure shows a Me-109G German fighter from World War II. Shown is the percentage breakdown of the drag (includes interference drag) of the components

Any physical body being propelled through the air has drag associated with it. In aerodynamics, drag is defined as the force that opposes forward motion through the atmosphere and is parallel to the direction of the free-stream velocity of the airflow. Drag must be overcome by thrust in order to achieve forward motion.

Drag is generated by nine conditions associated with the motion of air particles over the aircraft. There are several types of drag: form, pressure, skin friction, parasite, induced, and wave.

The term "separation" refers to the smooth flow of air as it closely hugs the surface of the wing then suddenly breaking free of the surface and creating a chaotic flow. The second picture on the left hand margin of this page shows examples of air flowing past a variety of objects. The bottom shows well behaved, laminar flow (flow in layers) where the flow stays attached (close to the surface) of the object. The object just above has a laminar flow for the first half of the object and then the flow begins to separate from the surface and form many chaotic tiny vortex flows called vortices. The two objects just above them have a large region of separated flow. The greater the region of separated flow, the greater the drag. This is why airplane designers go to such effort to streamline wings and tails and fuselages — to minimize drag.

### Induced drag

Induced drag is a by-product of lift

Induced drag is the drag created by the vortices at the tip of an aircraft's wing. Induced drag is the drag due to lift. The high pressure underneath the wing causes the airflow at the tips of the wings to curl around from bottom to top in a circular motion. This results in a trailing vortex. Induced drag increases in direct proportion to increases in the angle of attack. The circular motion creates a change in the angle of attack near the wing tip which causes an increase in drag. The greater the angle of attack up to the critical angle (where a stall takes place), the greater the amount of lift developed and the greater the induced drag.

All of these types of drag must be accounted for when determining drag for subsonic flight. The total drag is the sum of parasite and induced drag.

Total Drag = Parasite Drag + Induced Drag

But the net (or total) drag of an aircraft is not simply the sum of the drag of its components. When the components are combined into a complete aircraft, one component can affect the air flowing around and over the airplane, and hence, the drag of one component can affect the drag associated with another component. These effects are called interference effects, and the change in the sum of the component drags is called interference drag. Thus,

(Drag)1+2 = (Drag)1 + (Drag)2 + (Drag)interference

Generally, interference drag will add to the component drags but in a few cases, for example, adding tip tanks to a wing, total drag will be less than the sum of the two component drags because of the reduction of induced drag.

### parasite drag

The parasite drag of a typical airplane in the cruise configuration consists primarily of the skin friction, roughness, and pressure drag of the major components. There is usually some additional parasite drag due to such things as fuselage upsweep, control surface gaps, base areas, and other extraneous items. Since most of the elements that make up the total parasite drag are dependent on Reynolds number and since some are dependent on Mach number, it is necessary to specify the conditions under which the parasite drag is to be evaluated. In the method of these notes, the conditions selected are the Mach number and the Reynolds number corresponding to the flight condition of interest.

The basic parasite drag area for airfoil and body shapes can be computed from the following expression:
f = k cf Swet

where the skin friction coefficient, cf , which is based on the exposed wetted area includes the effects of roughness, and the form factor, k, accounts for the effects of both supervelocities and pressure drag. Swet is the total wetted area of the body or surface.

Computation of the overall parasite drag requires that we compute the drag area of each of the major components (fuselage, wing, nacelles and pylons, and tail surfaces) and then evaluate the additional parasite drag components described above.

We thus write:
CDp =
S ki cfi Sweti / Sref + CDupsweep + CDgap+ CDnac_base + CDmisc
where the first term includes skin friction, and pressure drag at zero lift of the major components. cfi is the average skin friction coefficient for a rough plate with transition at flight Reynolds number. Equivalent roughness is determined from flight test data.

### form drag

Form drag and pressure drag are virtually the same type of drag. Form or pressure drag is caused by the air that is flowing over the aircraft or airfoil. The separation of air creates turbulence and results in pockets of low and high pressure that leave a wake behind the airplane or airfoil (thus the name pressure drag). This opposes forward motion and is a component of the total drag. Since this drag is due to the shape, or form of the aircraft, it is also called form drag. Streamlining the aircraft will reduce form drag, and parts of an aircraft that do not lend themselves to streamlining are enclosed in covers called fairings, or a cowling for an engine, that have a streamlined shape. Airplane components that produce form drag include (1) the wing and wing flaps, (2) the fuselage, (3) tail surfaces, (4) nacelles, (5) landing gear, (6) wing tanks and external stores, and (7) engines.

Skin friction drag is caused by the actual contact of the air particles against the surface of the aircraft. This is the same as the friction between any two objects or substances. Because skin friction drag is an interaction between a solid (the airplane surface) and a gas (the air), the magnitude of skin friction drag depends on the properties of both the solid and the gas. For the solid airplane, skin fiction drag can be reduced, and airspeed can be increased somewhat, by keeping an aircraft's surface highly polished and clean. For the gas, the magnitude of the drag depends on the viscosity of the air. Along the solid surface of the airplane, a boundary layer of low energy flow is generated. The magnitude of the skin friction depends on the state of this flow.

### skin friction

The leading edge of a wing will always produce a certain amount of friction drag

An important aerodynamic force during low-speed subsonic flight is the shear force (the sideways force or internal friction) caused by viscous airflow over the surfaces of the vehicle. This shear force is referred to as the skin-friction force or skin-friction drag and depends strongly on the Reynolds number, surface roughness, and pressure gradients. In addition to the pressure forces that act everywhere perpendicular to (normal to) a body in moving air, viscous forces are also present. It is these viscous forces that modify the lift that would exist under ideal conditions (air is inviscid and incompressible) and help create the real drag.

If the airflow were ideal, that is, inviscid, the air would simply slip over the surface of a smooth plate with velocity V¥. At all points along the surface of the plate, the velocity distribution (that is, the variation in velocity as one moves perpendicularly away from the surface) would be a uniform constant value of V∞. No drag would result if the airflow were frictionless (inviscid).

Under real conditions, however, a very thin film of air molecules adheres to the surface. This is the very important no-slip condition. It states that at the surface of a body, the airflow velocity is zero. As one moves away from the body, the velocity of the air gradually increases until, at some point, the velocity becomes a constant value; in the case of a flat plate this value is V¥. The layer of air where the velocity is changing from zero to a constant value is called the boundary layer. Within the boundary layer, there are relative velocities between the layers and an internal friction is present. This internal friction extends to the surface of the body. The cumulative effect of all these friction forces is to produce drag on the plate. This drag is referred to as skin-friction drag.

Real fluid flow about an airfoil. The thickness of the boundary layers and wake are greatly exaggerated. The bottom flow along lower surface is the same as on the upper surface.

Initially, near the leading edge of a flat, smooth plate, one has a laminar flow (the flow is layered) and the boundary layer also is steady and layered—hence, a laminar boundary layer. As one moves farther downstream, viscosity continues to act, and the laminar boundary layer thickens as more and more air is slowed down by internal friction. Eventually, a point is reached on the plate where the laminar boundary layer undergoes transition and becomes a turbulent boundary layer. As is usual for turbulent flow, there is random motion in the boundary layer as well as the downstream-directed motion. There is no slip at the surface of the plate. Another important difference from the laminar boundary layer is the fact that the velocity builds up more quickly as one moves away from the wall, although the total boundary-layer thickness is greater. The turbulent boundary layer farther away from the wall reenergizes the slower moving air nearer the wall. This condition can be seen by comparing the profile of the laminar boundary layer with the profile of the turbulent boundary layer.

The Reynolds number has an important effect on the boundary layer. As the Reynolds number increases (caused by increasing the airflow speed and/or decreasing the viscosity), the boundary layer thickens more slowly. However, even though the Reynolds number becomes large, the velocity at the surface of the body must be zero. Thus, the boundary layer never disappears.

It is interesting to note that a typical thickness of the boundary layer on an aircraft wing is generally less than a centimetre (2.5 inches). Yet, the velocity must vary from zero at the surface of the wing to hundreds of meters per second at the outer edge of the boundary layer. It is evident that tremendous shearing forces (internal friction) must be acting in this region. This gives rise to the skin-friction drag.

Applied to an airfoil in a real airflow, the same free-stream velocity V¥ and free-stream static pressure p¥ apply. The field of air ahead of the airfoil is only slightly modified and for all practical purposes, the velocities and static pressures are the same as for the ideal fluid case. Again a stagnation point (a point with no motion) occurs at the leading edge of the airfoil and the pressure reaches its maximum value of pt at this point (total or stagnation pressure). From this point on along the airfoil, the picture changes.

As noted earlier in the example of the flat plate, a boundary layer begins to form because of viscosity. This boundary layer is very thin and outside of it, the flow acts very much like that of an ideal fluid. Also, the static pressure acting on the surface of the airfoil is determined by the static pressure outside the boundary layer. This pressure is transmitted through the boundary layer to the surface and thus acts as if the boundary layer were not present at all. But the boundary layer feels this static pressure and will respond to it.

Over the front surface of the airfoil up to the shoulder, an assisting favourable pressure gradient exists (pressure decreasing with distance downstream). The airflow speeds up along the airfoil. The flow is laminar and a laminar boundary layer is present. This laminar boundary layer grows in thickness along the airfoil. When the shoulder is reached, however, the air molecules are moving slower than in the ideal fluid case. This is an unfavourable condition because the previous ideal flow just came to rest at the trailing edge of the airfoil. It would appear now, with viscosity present, that the flow will come to rest at some distance before the trailing edge is reached.

As the airflow moves from the shoulder to the rear surface of the airfoil, the static-pressure gradient is unfavourable (increasing pressure with downstream distance). The air molecules must push against both this unfavourable pressure gradient and the viscous forces. At the transition point, the character of the airflow changes and the laminar boundary layer quickly becomes a turbulent boundary layer. This turbulent boundary layer continues to thicken downstream. Pushing against an unfavourable pressure gradient and viscosity is too much for the airflow, and at some point, the airflow stops completely. The boundary layer has stalled short of reaching the trailing edge. (Remember that the airflow reached the trailing edge before stopping in the ideal fluid case.)

This stall point is known as the separation point. All along a line starting from this point outward into the airflow, the airflow is stalling. Beyond this line, the airflow is actually moving backward, upstream toward the nose before turning around. This is a region of eddies and whirlpools and represents “dead” air that is disrupting the flow field away from the airfoil. Thus, the airflow outside the dead air region is forced to flow away and around it. The region of eddies is called the wake behind the airfoil.

Up to the separation point, the difference between the static-pressure distribution for ideal fluid flow and real airflow is not very large but once separation occurs, the pressure field in greatly modified. In the ideal fluid case, the net static-pressure force acting on the front surface of the airfoil (up to the shoulder) parallel to the free stream exactly opposed and cancelled that acting on the rear surfaces of the airfoil. Under real airflow conditions, however, this symmetry and cancellation of forces is destroyed. The net static-pressure force acting on the front surface parallel to the free-stream direction now exceeds that acting on the rear surface. The net result is a drag force due to the asymmetric pressure distribution called pressure drag. This is a drag in addition to the skin-friction drag due to the shearing forces (internal friction) in the boundary layer. Additionally, the modification of the static-pressure distribution causes a decrease in the pressure lift from the ideal fluid case. The effect of viscosity is that the lift is reduced and a total drag composed of skin-friction drag and pressure drag is present. Both of these are detrimental effects.

It should be emphasized that similar processes are occurring on all the components of the aircraft to one degree or another, not only the airfoil.

Thus, the effects of a real fluid flow are the result of the viscosity of the fluid. The viscosity causes a boundary layer and, hence, a skin-friction drag. The flow field is disrupted because of viscosity to the extent that a pressure drag arises. Also, the net pressure lift is reduced.

### interference drag

Surfaces at angles to each other as in (C) create turbulence in the region of the joint. This occurs most frequently at the intersection of the fuselage and wing.

This figure shows a Grumman F9F Panther Jet with a large degree of filleting to reduce drag

### lift versus drag

An aircraft with a given total gross weight can be operated in level flight over a range of power settings and airspeeds. Since Lift and Weight must be equal in order to maintain level flight, it is obvious that there is a relationship between Lift (L), Airspeed (V), and Angle of Attack (AT). This relationship can be "generalized" with the following expression. (Note: the expression is not an exact equation).

Lift = Angle of Attack x Velocity

Since angle of attack and speed also have a relationship to Induced Drag and Parasite Drag, the relationship of Lift/Drag is shown by the graph below.

Parasite drag increases with speed. Induced drag decreases with speed. The SUM of the two drags (Total Drag curve) shows that there is only one airspeed for a given airplane and load that provides MINIMUM total drag. This is the point M which is the maximum lift over drag ratio (L/D). It is the airspeed at which the aircraft can glide the farthest without power (maximum glide range). This is the airspeed which should immediately be set up in the event of a power failure. This maximum glide airspeed is different for each aircraft design. The Pilot Operating Handbook should be consulted for this airspeed and the pilot should memorize it to eliminate need to search manuals during an emergency.

Decrease in airplane drag coefficient with time

 Your browser does not support inline frames or is currently configured not to display inline frames.