Airfoils and Lift

The other critical feature present in both aircraft was the use of a plywood leading-edge contour panel. This was relatively unusual in WWI. British aircraft seem not to have used it at all, preferring intermediate riblets as leading-edge support; and from a quick appraisal of my library, I have identified only 5 aircraft which had this feature (I don't suppose this to be at a definitive.). These are the Pfalz D.XII Fokkers Dr.I, D.VI and D.VII, and the Nieuport 28 (possibly also the 27).

Some aircraft wings were, of course, totally skinned in sheet plywood or aluminum; but with these exceptions, at least, complete fabric cover was the norm. The use of plywood leading-edge covering presents a problem in the attachment of fabric since stringing (ie, the through-wing stitching normally used) would be required to stop at the plywood-covered surface. This may account for the fact that both the triplane and N28 are reported as originally having the fabric tacked to the rib flanges rather than being sewn (which was considered to be the correct way). The fabric attachment itself is therefore suspect but the test still remains; why only failures of the upper wing? If the fabric attachment was the critical factor, then failures could have occurred in any wing with this feature, which would have included lower planes of both the triplane and the N28.

Both aircraft have structurally suspect features in their wing leading-edges. In the case of the N28, the long rib-noses would produce large bending stresses (during violent manoeuvres) at their main-spar attachment locations. Large bending stresses can have attendant large shear stresses; and on the N28, these would exist in the thin poplar rib-webs (typical of the period). This is a very risky arrangement, since timber is not particularly strong when subject to shear loading along the grain - plywood is much better. (The N28 rib-noses had very little shear material anyway)

The other suspect feature is that of the omission of rib-capping referred to in the recent WWI AERO article. These details appear peculiar to the N28, and are at the most extreme in the upper wing. There is little doubt that the upper wing leading edge was simply of marginal strength; and at first sight it seems odd that sandloading did not reveal this weakness. But of course this reveals a weakness of sand-loading. The chordwise distribution of lift, at high angles of attack, will not normally be represented by a heap of sand, since dry sand slumps to approximately 45 deg- forming a triangular load distribution with a centrally-located center of gravity. (This can be modified within limits by constructing walls along the wing edges.) Sandloading therefore successfully tests the wingspar adequacy, but is insufficient to the task of testing the rib nose strength (and remember that here we have 2 aircraft which resolutely held on to their spars, whilst liberally shedding secondary structure). This proof-loading problem is exacerbated by the fact that wing lift (particularly at large angles of attack) is largely generated by the negative pressure zone existing on the forward upper surface (see Fig 18- taken from SIMPLE AERODYNAMICS (1929), by Charles N Monteith.).

The critical structural requirement under these loading conditions is to have adequate "peel" strength between the upper skin and the substructure (ribs and/or stringers etc). Both the N28 and the Dr.I were deficient here. The Nieuport was devoid of rib cap-strips or spanwise stringers at the critical location; the Dr.I leading-edge plywood was severely cut away at each rib, had no supporting stringers, and had only minor connection to the main spar. With this arrangement, a significant amount of the local lift- would have been transmitted in a peel condition from the plywood skin to the supporting ribs - there was no other load-path. Again, this is a very unreliable form of joint. Today, the attachment of wing skins to substructure remains a critical factor; in fact, where fuel is carried inside a wing much of the wing design is overridingly determined by this consideration.

So, the Nieuport had a weak upper-wing leading edge and larger chord to boot. This could (as suggested in the WWI AERO article) be the complete answer to the N28 failures. But the Triplane had the same design condition on all wings, but only the top wing ever failed. So there was something else.

It is not common to see a biplane or triplane wing cellule in which equal-chord wings are of differing span, although some famous aircraft such as the BE- 12, RE-8 and Curtiss Jenny are exceptions. Typically, where an upper wing is of greater span, it is often of greater chord also. This has the virtue of approximately maintaining constant aspect-ratio for each wing in the complete wing system. (To what extent this represented a design objective at the time I have no information.)

The fact that real wings are of finite span (as opposed to the theoretically infinite span wing which is implicit in aerofoil section data) means that a real wing will attain a particular lift coefficient at an angle of attack somewhat greater than that apparent from he section-data. It also follows that wings of differing aspect ratio, but identical section, will generate different lift-intensities, to one another, when operating at the same angle of attack.

The Dr.1 had aspect ratios of 6.8, 5.9 and 5.1 for the upper, middle and lower planes respectively. The wing section (tested as the Gottingen 289 section after the war) had a maximum lift coefficient of about 1.4. Making estimates for each of the triplane wings (working as independent surfaces), the planes would require 19.2, 20 and 21 degrees respectively to reach the maximum lift coefficient. When working at the same angle of attack (as in the aircraft alignment), the upper wing would produce a lift intensity about 9% greater than the lower wing. So could aspect-ratio be the cause of the Triplane wing failures? Well no, I am afraid not. A 9% increased lift intensity cannot be considered sufficient to always fail the upper wing before one or the other planes. Variations in material strength and build quality would both have similar (or greater) tolerance, which would occasionally bias the failure to one of the other planes. There has to be something else - something more emphatic.

I found the answer by chance, and I found it in a `history' book. Whilst flipping through a copy of SIMPLE AERODYNAMICS (1929), by Charles N Monteith, (Chief Engineer, Boeing), looking for data on the Gottingen 289 section, I came across a particularly relevant passage under Item 70, p89, “Pressure distribution tests on MB-3A Airplane”, which is reproduced in facsimile here:

Paragraphs B and C are telling. The loading distribution noted is very significant over the biplane system described. A factor of 1.6 at high-lift coefficients cannot be ignored. The Triplane system with its relatively smaller wing gaps and pronounced stagger would almost certainly have a greater value than this. Together with aspect-ratio effects it is not unreasonable to suggest that the lift intensity of the upper wing of the Dr.I approached twice that of the bottom wing. This is certainly enough to test the upper wing integrity before the rest of the system.

Conclusion

I would suggest that the Dr.I wing failures (and almost certainly those of the N28, too) occurred because lift-grading (particularly), together with aspect-ratio effects, caused the upper surface of the upper wing to be subject to much greater lift intensity than the rest of the system. This tested a leading-edge design of marginal strength, poorly made, to the point of collapse in particular aircraft. The leading edge failure continued back across the wing due to design details. Where rib tails, for example, were connected by a wire trailing edge, ballooning fabric will exert tensile loading in this wire which will then tend to "gather up" the rib tails and strip the wing. This would also destabilize the area of the aileron support structures, and so on. The strengthening of the wing aft of the spars and the improvements to build quality, carried out after the original failures, would have acted to prevent this catastrophic failure. But the root cause of the failure lift-grading) went unappreciated until after the war when investigations like those at NACA were conducted.

It would be fascinating to know to what extent these factors were understood prior to 1918. I expect that the concentration of lift forces (as an intense negative pressure zone at the upper surface LE) was reasonably well appreciated by wind-tunnel investigators- if only by the application of Bernoulli's theorem to the visible flow patterns around test sections. Probably the effects of aspect ratio were understood- even if only qualitatively; but lift-grading would require much more complex investigation. Regarding the aspect-ratio issue; advocates of multiplanes (Horatio Phillips, for example) appear to have worked from the understanding that high aspect-ratio is a "good thing" (true) but not to have had evidence of the detrimental effects of interference between closely-spaced multi-plane wing systems.

But such is the nature of progress - the testing of ideas. It took the lives of airmen to drive the investigations which led to today's understanding of these matters and which allow our complacent and sometimes arrogant review of history.

A final thought. It is theoretically possible for the Fokker triplane to remain airborne on its 2 lower planes alone (of 9.9 square metres area). The stall speed would be about 64mph. No doubt, when both Gontermann and Pastor found themselves in dire straits, they did the natural thing: to pull back on the stick even though the aircraft was deeply stalled. Maybe if they had first pushed... ?

Forces Acting on an Airplane

Lift opposes gravity.

Thrust opposes drag.

Drag and weight are forces inherent in anything lifted from the earth and moved through the air. Thrust and lift are artificially created forces used to overcome the forces of nature and enable an airplane to fly. The engine and propeller combination is designed to produce thrust to overcome drag. The wing is designed to produce lift to overcome the weight (or gravity).

In straight-and-level, unaccelerated flight, (Straight-and-level flight is coordinated flight at a constant altitude and heading) lift equals weight and thrust equals drag, though lift and weight will not equal thrust and drag. Any inequality between lift and weight will result in the airplane entering a climb or descent. Any inequality between thrust and drag while maintaining straight-and-level flight will result in acceleration or deceleration until the two forces become balanced.

Flight Control Surfaces

Laminar Flow is the smooth, uninterrupted flow of air over the contour of the wings, fuselage, or other parts of an aircraft in flight. Laminar flow is most often found at the front of a streamlined body and is an important factor in flight. If the smooth flow of air is interrupted over a wing section, turbulence is created which results in a loss of lift and a high degree of drag. An airfoil designed for minimum drag and uninterrupted flow of the boundary layer is called a laminar airfoil.

The Laminar flow theory dealt with the development of a symmetrical airfoil section which had the same curvature on both the upper and lower surface. The design was relatively thin at the leading edge and progressively widened to a point of greatest thickness as far aft as possible. The theory in using an airfoil of this design was to maintain the adhesion of the boundary layers of airflow which are present in flight as far aft of the leading edge as possible. on normal airfoils the boundary layer would be interrupted at high speeds and the resultant break would cause a turbulent flow over the remainder of the foil. This turbulence would be realized as drag up the point of maximum speed at which time the control surfaces and aircraft flying characteristics would be affected. The formation of the boundary layer is a process of layers of air formed one next to the other, ie; the term laminar is derived from the lamination principle involved.

History of Laminar Flow The P-51 Mustang is the first aircraft every intentionally designed to use laminar flow airfoils. However, wartime NACA research data I have shows that Mustangs were not manufactured with a sufficient degree of surface quality to maintain much laminar flow on the wing. The RAE found that the P-63, despite being designed with laminar airfoils, also was not manufactured with sufficient surface quality to have much laminar flow.

The Mustang was a mathematically designed airplane and the wing foil that was to be classified as a "semi-empirical venture" by the British was cleared for adoption on the new design. The wing section would be the only part of the fighter which would be tested in a wind tunnel prior to the first test flight. Due to the speculation of the success of the radical foil, the engineering department was committed to adopt a more conventional airfoil within thirty days of the tests in the event the wing did not come up to specifications. A one quarter scale model of the wing was designed and constructed for tests in the wing tunnel at the Caiifornia Institute of Technology.

The use of this airfoil on the Mustang would greatly add to the drag reducing concept that was paramount in all design phases of the airplane. The few applications of this foil, prior to this time, had been handbuilt structures which were finished to exacting tolerances. An absolutely smooth surface was necessary due to the fact that any surface break or rough protrusion would interrupt the airflow and detract from the laminar flow theory. Because of the exactness required, the foil had been shelved by other manufacturers due to the clearances and tolerances which are used in mass production. The engineers at NAA approached this problem with a plan to fill and paint the wing surface to provide the necessary smoothness. The foil which was used for the Mustang had a thickness ratio of 15.1 percent at the wing root at 39 percent of the chord. The tip ratio was 11.4 percent at the 50 percent chord line. These figures provided the maximum thickness area at 40 percent from the leading edge of the wing and resulted in a small negative pressure gradient over the leading 50-60 percent of the wing surface.

The B-24 bomber's "Davis" airfoil was also a laminar flow airfoil, which predates the Mustang's. However, the designers of the B-24 only knew that their airfoil had very low drag in the wind tunnel. They did not know that it was a laminar flow airfoil.

There were several aircraft modified by NACA, in the late 1930s, to have laminar flow test sections on their wings. Hence, such aircraft as a modified B-18 were some of the first aircraft to fly with laminar flow airfoils.

The boundary layer concept is credited to the great German aerodynamicist, Ludwig Prandtl. Prandtl hyposthesized and proved the existence of the boundary layer long before the Mustang was a gleam in anyone's eye.

Example: First, lets get more specific about what laminar flow is. The flow next to any surface forms a "boundary layer", as the flow has zero velocity right at the surface and some distance out from the surface it flows at the same velocity as the local "outside" flow. If this boundary layer flows in parllel layers, with no energy transfer between layers, it is laminar. If there is energy transfer, it is turbulent.

All boundary layers start off as laminar. Many influences can act to destabilize a laminar boundary layer, causing it to transition to turbulent. Adverse pressure gradients, surface roughness, heat and acoustic energy all examples of destabilizing influences. Once the boundary layer transitions, the skin friction goes up. This is the primary result of a turbulent boundary layer. The old "lift loss" myth is just that - a myth.

A favorable pressure gradient is required to maintain laminar flow. Laminar flow airfoils are designed to have long favorable pressure gradients. All airfoils must have adverse pressure gradients on their aft end. The usual definition of a laminar flow airfoil is that the favorable pressure gradient ends somewhere between 30 and 75% of chord.

Now Consider the finish on your car in non-rainy conditions. Dust and leaves have settled on the hood's paint. We go for a drive. At once the leaves blow off. But the dust remains. We speed up. Even if we go very fast, the dust remains because of the thin layer of air that moves with the car. If you drive with dew on your car, the dew will not so quickly be blown dry where the air flow has this thin laminar layer. Downstream, where the laminar flow has become turbulent, the air flow quickly dries the dew.

In the fifties this was dramatically shown in a photograph of the top of a sailplane wing (inflight) that had dew on it. A few tiny seeds had landed on forward area the wing while on the ground. In flight these seeds, tiny though they were, reached through the laminar layer and caused micro-turbulence causing the dew to be blown dried in an expanding vee shaped area down stream of each tiny seed.

Additional information

Profile drag

This comprises two components: surface friction drag and normal pressure drag (form drag).

Surface friction drag

This arises from the tangential stresses due to the viscosity or "stickiness" of the air. When air flows over any part of an aircraft there exists, immediately adjacent to the surface, a thin layer of air called the boundary layer, within which the air slows from its high velocity at the edge of the layer to a standstill at the surface itself. Surface friction drag depends upon the rate of change of velocity through the boundary layer, i.e. the velocity gradient. There are two types of boundary layer, laminar and turbulent, the essential features of which are shown in Fig 8. Although all combat aircraft surfaces develop a laminar boundary layer to start with, this rapidly becomes turbulent within a few per cent of the length of the surface. This leaves most of the aircraft immersed in a turbulent boundary layer, the thickness of which increases with length along the surface. The velocity and hence pressure variations along the length of any surface can have adverse effects on the behavior of the boundary layer, as will be discussed later.

Surface friction drag can amount to more than 30% of the total drag under cruise conditions.

Normal pressure drag (form drag)

This also depends upon the viscosity of the air and is related to flow separation. It is best explained by considering a typical pressure distribution over a wing section, as shown in Fig 4, first at low AOA and then at high AOA.

At low AOA the high pressures near the leading edge produce a component of force in the rearward (i.e. drag) direction, while the low pressures ahead of the maximum thickness point tend to suck the wing section forward, giving a thrust effect. The low pressures aft of the maximum thickness point tend to suck the wing rearwards, since they act on rearward-facing surfaces. Without the influence of the boundary layer, the normal pressure forces due to the above drag and thrust components would exactly cancel.

There is a favorable pressure gradient up to the minimum pressure point, with the pressure falling in the direction of flow. This helps to stabilize the boundary layer. Downstream of the minimum pressure point, however, the thickening boundary layer has to flow against an adverse pressure gradient. Viscous effects reduce momentum within the boundary layer, and the thickness of the layer further increases so that the external flow "sees" a body which does not appear to close to a point at the trailing edge. A narrow wake is formed as the boundary layer streams off the section. This prevents the pressures on the aft-facing surface of the wing section from recovering to the high value obtaining near the stagnation point on the leading edge, as they would have done if a boundary layer had not formed. There is thus a lower than expected pressure acting on the aftfacing surface, giving rise to normal pressure drag. In the low-AOA case this component is small, most of the profile drag being made up of surface friction drag.

As the AOA of the wing section is increased, the point of minimum pressure moves towards the leading edge, with increasingly high suction being achieved. This means that the pressure then has to rise by a greater extent downstream of the minimum pressure point and that the length of wing surface exposed to the rising pressure is increased. The resulting adverse pressure gradient becomes more severe as AOA is increased. This has serious implications for the boundary layer, which is always likely to separate from the wing surface under such conditions.

The Swept Wing

The whole idea of sweeping an aircraft's wing is to delay the drag rise caused by the formation of shock waves. The swept-wing concept had been appreciated by German aerodynamicists since the mid-1930s, and by 1942 a considerable amount of research had gone into it. However, in the United States and Great Britain, the concept of the swept wing remained virtually unknown until the end of the war. Due to the early research in this area, this allowed Germany to successfully introduce the swept wing in the jet fighter Messerschmitt ME-262 as early as 1941.

Early British and American jet aircraft were therefore of conventional straight-wing design, with a high-speed performance that was consequently limited. Such aircraft included the UKGloster Meteor F.4 , the U.S. Lockheed F-80 Sooting Star and the experimental U.S. jet, the Bell XP-59A Airacomet.

After the war German advanced aeronautical research data became available to the United States Army Air Force (USAAF) as well as Great Britain. This technology was then incorporated into their aircraft designs. Some early jets that took advantage of this technology were the North American F-86 Sabre, the Hawker Hunter F.4 and the Supermarine Swift FR.5.

Not to be outdone, the Soviet Union introduced the swept wing in the Mikoyan Mig-15 in 1947. This aircraft was the great rival of the North American F-86 Sabre during the Korean War.

Jet Engine Theory

Centuries ago in 100 A.D., Hero, a Greek philosopher and mathematician, demonstrated jet power in a machine called an "aeolipile." A heated, water filled steel ball with nozzles spun as steam escaped.

Over the course of the past half a century, jet-powered flight has vastly changed the way we all live. However, the basic principle of jet propulsion is neither new nor complicated.

Centuries ago in 100 A.D., Hero, a Greek philosopher and mathematician, demonstrated jet power in a machine called an "aeolipile." A heated, water filled steel ball with nozzles spun as steam escaped. Why? The principle behind this phenomenon was not fully understood until 1690 A.D. when Sir Isaac Newton in England formulated the principle of Hero's jet propulsion "aeolipile" in scientific terms. His Third Law of Motion stated: "Every action produces a reaction... equal in force and opposite in direction."

The jet engine of today operates according to this same basic principle. Jet engines contain three common components: the compressor, the combustor, and the turbine. To this basic engine, other components may be added, including:

· A nozzle to recover and direct the gas energy and possibly divert the thrust for vertical takeoff and landing as well as changing direction of aircraft flight.

· An afterburneror augmentor, a long "tailpipe" behind the turbine into which additional fuel is sprayed and burned to provide additional thrust.

· A thrust reverser, which blocks the gas rushing toward the rear of the engine, thus forcing the gases forward to provide additional braking of aircraft.

· A fan in front of the compressor to increase thrust and reduce fuel consumption.

· An additional turbine that can be utilized to drive a propeller or helicopter rotor.

The Turbojet Engine

A turbojet engine.

The turbojet is the basic engine of the jet age. Air is drawn into the engine through the front intake. The compressor squeezes the air to many times normal atmospheric pressure and forces it into the combustor. Here, fuel is sprayed into the compressed air, is ignited and burned continuously like a blowtorch. The burning gases expand rapidly rearward and pass through the turbine. The turbine extracts energy from the expanding gases to drive the compressor, which intakes more air. After leaving the turbine, the hot gases exit at the rear of the engine, giving the aircraft its forward push... action, reaction !

For additional thrust or power, an afterburner or augmentor can be added. Additional fuel is introduced into the hot exhaust and burned with a resultant increase of up to 50 percent in engine thrust by way of even higher velocity and more push.

The Turboprop/Turboshaft Engine

A turboprop, or turboshaft engine.

A turboprop engine uses thrust to turn a propeller. As in a turbojet, hot gases flowing through the engine rotate a turbine wheel that drives the compressor. The gases then pass through another turbine, called a power turbine. This power turbine is coupled to the shaft, which drives the propeller through gear connections.

A turboshaft is similar to a turboprop engine, differing primarily in the function of the turbine shaft. Instead of driving a propeller, the turbine shaft is connected to a transmission system that drives helicopter rotor blades; electrical generators, compressors and pumps; and marine propulsion drives for naval vessels, cargo ships, high speed passenger ships, hydrofoils and other vessels.

The Turbofan Engine

A high bypass turbofan engine.

A turbofan engine is basically a turbojet to which a fan has been added. Large fans can be placed at either the front or rear of the engine to create high bypass ratios for subsonic flight. In the case of a front fan, the fan is driven by a second turbine, located behind the primary turbine that drives the main compressor. The fan causes more air to flow around (bypass) the engine. This produces greater thrust and reduces specific fuel consumption.

A low bypass turbofan engine.

For supersonic flight, a low bypass fan is utilized, and an augmentor is added for additional thrust.