Inventor of the foreplane Gustav Lohmann proposed at the end of the 1930s to equip a tailless aircraft with a freely floating wing positioned in front of the main wing. This wing was equipped with a servo rudder, which adjusted its lift. It served to compensate for the additional pitching moment of the wing that occurs during flap deployment. Since Lohmann was an employee of Handley Page, the company owned the patent for this technical solution, and this idea is mentioned in the technical literature under that brand. However, there has been no practical embodiment of this idea so far! What is the reason?
Losses due to balancing
The wing of an aircraft that generates lift has a concurrent, we might say, negative byproduct in the form of a pitching moment that tends to pitch the aircraft downwards. To prevent the aircraft from pitching, a small wingâthe stabilizerâexists at its tail, which counters this pitching by creating a downward, or negative, lift. This aerodynamic scheme of an aircraft is called "normal." Since the lift of the stabilizer is negative, it adds to the weight of the aircraft, necessitating that the wing generates lift greater than the weight.
The difference between these forces is referred to as balancing losses, which can reach up to 20%.
However, the first flying airplane by the Wright Brothers did not experience such losses because the small wingâa destabilizer that prevents pitchingâwas located not behind the wing but in front of it. This aerodynamic scheme of an aircraft is called "duck." To prevent the aircraft from pitching, the destabilizer must generate lift directed upwards, or positive. This lift adds to the lift from the wing, and this sum equates to the weight of the aircraft. As a result, the wing must create lift that is less than the weight. Thus, there are no balancing losses!
The stabilizer and destabilizer are collectively referred to as horizontal tail surfaces or HT.
However, with the widespread development in the early 1930s of wing lift augmentation systems, the "duck" lost this advantage. The primary element of such systems is the flapâthe hinged rear part of the wing that deflects downward. It roughly doubles the lift generated by the wing, allowing for reduced takeoff and landing speeds, thereby saving on landing gear mass. However, the side effect is a significant increase in the pitching moment when deploying the flap, to the extent that the destabilizer cannot manage it, while the stabilizer can. Breaking is not the same as building; in this case, that is a positive force.
For a wing to generate lift, it must be oriented at an angle to the oncoming airflow. This angle is known as the angle of attack, and as it increases, so does the lift, but not indefinitelyâonly up to a critical angle, which ranges from 15 to 25 degrees. Consequently, the total aerodynamic force is directed not strictly upward but tilts toward the tail of the aircraft. This force can be decomposed into a component directed strictly upwardâliftâand a component directed backwardâdrag. The ratio of lift to drag indicates the aerodynamic quality of the aircraft, which can range from 7 to 25.
A phenomenon supporting the conventional scheme is the airflow skew behind the wing, which involves a downward deflection of the airflow that increases with the lift produced by the wing. Therefore, when the flap is deflected, the actual negative angle of attack of the stabilizer increases automatically due to aerodynamics, leading to a corresponding negative lift.
In addition, the advantage of the "normal" scheme over the "duck" scheme is supported by the consideration of longitudinal flight stability of the aircraft. The angle of attack of the aircraft can change due to vertical movements of air masses. Aircraft are designed with this phenomenon in mind and aim to counteract disturbances. Each surface of the aircraft has an aerodynamic focus â a point of application for the increment in lift force when the angle of attack changes. If we consider the resultant increments from the wing and the horizontal stabilizer, there is also a focus for the aircraft. If the focus of the aircraft is behind the center of mass, any accidental increase in the angle of attack will cause the increase in lift to tilt the aircraft in such a way as to decrease the angle of attack. The aircraft then returns to its previous flight mode. In this case, in the "normal" configuration, the wing creates a destabilizing moment (with an increase in angle of attack), while the stabilizer creates a stabilizing moment (with a decrease in angle of attack), and the latter prevails by about 10%. In the "duck" configuration, the destabilizing moment is generated by the destabilizer, while the stabilizing moment is generated by the wing, which is approximately 10% greater. Therefore, increasing the area and leverage of the horizontal stabilizer leads to increased stability in the normal configuration and decreased stability in the "duck" configuration. All moments act and are considered relative to the center of mass of the aircraft (see Fig. 1).
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If the center of gravity of the aircraft is ahead of its center of mass, then with a slight increase in the angle of attack, it increases even more, and the aircraft will be statically unstable. This configuration of the center of gravity and center of mass is used in modern fighters to load the stabilizer and achieve not negative, but positive lift. The flight of the aircraft is maintained not by aerodynamics, but by a quadruple redundant automated artificial stability system that 'steers' when the aircraft deviates from the required angle of attack. When the automation is turned off, the aircraft begins to pitch forward, which forms the basis for the 'Pugachev Cobra' maneuver, in which the pilot deliberately turns off the automation and upon reaching the required angle of tail pitch releases a missile into the rear hemisphere before re-engaging the automation.
In the future, we will only consider statically stable aircraft, as only such aircraft can be used in civil aviation.
The mutual positioning of the aircraft's center of gravity and center of mass is characterized by the term 'balance'.
Since the center of gravity is located behind the center of mass regardless of the design, the distance between them, known as the stability margin, increases the lever arm in a normal configuration and decreases it in a 'duck'.
The ratio of wing span to arm in a 'duck' is such that the lift force of the destabilizer at maximum deflection of the elevators is fully utilized when bringing the aircraft to high angles of attack. And it will not be sufficient when the flaps are deployed. This is why all 'ducks' designed by the famous American engineer Rutan lack any form of mechanization. His aircraft, the 'Voyager', was the first in the world to fly around the globe without landing and refueling in 1986.
The exception is the Beechcraft 'Starship', but there, in order to use flaps, a very complex design with variable geometry of the destabilizer was employed, which was not brought to a serially reproducible state, leading to the project's closure.
The arm of the wing largely depends on how much the lift force of the destabilizer increases with a one-degree increase in its angle of attack; this parameter is called the derivative of the lift coefficient with respect to angle of attack or simply the destabilizer's derivative. And the smaller this derivative is, the closer the center of mass of the aircraft can be placed to the wing, and thus, the shorter the wing arm will be. To reduce this derivative, the author proposed in 1992 to construct the destabilizer using a biplane scheme (2). This allows for such a reduction in the wing arm that it eliminates the obstacle to using a flap on it. However, there is a side effect of increased drag from the biplane configuration. Additionally, the aircraft's design is complicated, as it requires the construction of effectively two destabilizers instead of one.
Colleagues pointed out that the "biplane destabilizer" feature is present on the Wright Brothers' aircraft, but patents cover not only new features but also new combinations of features. The Wrights did not have a "flap" feature. Furthermore, if the combination of features of a new invention is known, at least one feature must be used for new purposes for the invention to be recognized. The Wrights used the biplane configuration to reduce the weight of the structure, while in the described invention, it is used to reduce the derivative.
Weathercock Duck
Almost two decades ago, the idea of the "weathercock duck" was recalled, mentioned at the beginning of the article.
It utilizes a weathercock horizontal stabilizer as a destabilizer, consisting of the stabilizer itself, hinged on an axis perpendicular to the fuselage, and a servo rudder connected to the stabilizer. Essentially, it's a small aircraft of a conventional layout, where the wing of the aircraft is the destabilizer of the weathercock, and the stabilizer of the aircraft is the servo rudder of the weathercock. This small aircraft doesn't fly; rather, it is positioned on an axis and self-orients itself concerning the oncoming flow. By changing the negative angle of attack of the servo rudder, we alter the angle of attack of the destabilizer relative to the flow, and, consequently, the lift of the weathercock when controlling pitch.
With the position of the control surface unchanged relative to the destabilizer, the FGO does not react to vertical wind gusts, i.e., changes in the aircraft's angle of attack. Therefore, its derivative equals zero. Based on our previous discussions, this is the ideal scenario.
During the testing of the first aircraft of the 'weather vane duck' design by A. Yurkounenko (3) with the FGO effectively loaded, more than twenty successful approaches were carried out. At the same time, clear signs of aircraft instability were observed (4).
Super-stability
Paradoxically, the instability of the 'weather vane duck' is a consequence of its 'super-stability'. The stabilizing moment of a classical duck with a fixed center of gravity is formed from the stabilizing moment of the wing and the opposing destabilizing moment at the center of gravity. In the weather vane duck, the FGO does not participate in forming the stabilizing moment, which is generated solely from the stabilizing moment of the wing. Thus, the stabilizing moment in the 'weather vane duck' is approximately ten times greater than that of a classical duck. With a random increase in the angle of attack, the aircraft, due to an excessive stabilizing moment from the wing, does not return to its previous state but 'skips' it. After the 'skipping', the aircraft has a reduced angle of attack compared to its previous state, which creates a stabilizing moment of the opposite sign, also excessive, resulting in oscillations that the pilot cannot dampen.
One condition for stability is the aircraft's ability to mitigate the effects of atmospheric disturbances. Therefore, in the absence of disturbances, satisfactory flight of an unstable aircraft is possible. This explains the successful approaches of the YUAN-1 aircraft. In the distant past, the author encountered a case where a new model glider flew for a total of at least 45 minutes in calm evening conditions, demonstrating quite satisfactory flights while exhibiting clear instabilityâpitching alternated with nosediving in the very first flight in windy weather. As long as the weather was calm and there were no disturbances, the glider showed satisfactory flight performance, but its adjustments were unstable. There were simply no reasons to exhibit this instability.
The described FGO can essentially be used in a "pseudo-duck". This type of aircraft is essentially a configuration of a "tailless" design and has an appropriate center of gravity. The FGO is only used to compensate for the additional pitching moment of the wing that occurs when the flaps are deployed. In cruising configuration, there is no load on the FGO. Thus, in its main operational flight mode, the FGO is essentially inactive, making its use in this configuration rather unproductive.
"KRASNOV-DUCK"
"Super Stability" can be eliminated by increasing the FGO derivative from zero to an acceptable level. This objective is achieved because the angle of rotation of the FGO is significantly less than the angle of rotation of the servo control caused by the change in the angle of attack of the aircraft (5). A rather simple mechanism depicted in Fig. 2 serves this purpose. FGO 1 and servo control 3 are pivotally mounted on axis O01. Rods 4 and 6 connect FGO 1 and servo control 3 to lever 8 through hinges 5, 7, 9, and 10. Coupling 12 is used for the pilot to adjust the length of rod 6 for pitch control. The rotation of FGO 1 occurs not over the full deflection angle of the servo control 3 relative to the aircraft when changing the direction of oncoming airflow, but only over its proportional part. If the ratio is half, then when an updraft increases the angle of attack of the aircraft by 2 degrees, the actual angle of attack of the FGO will increase by just 1 degree. Accordingly, the FGO derivative will be half as much compared to a fixed CG. The dashed lines indicate the positions of FGO 1 and servo control 3 after the change in the angle of attack of the aircraft. Changing the ratio and thus determining the magnitude of the derivative can easily be achieved by selecting the appropriate distances of hinges 5 and 7 from axis O01.
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Reducing the derivative of the CG through flapping allows for the placement of the focus and thereby the center of mass of the aircraft within any limits. This is the concept of aerodynamic offset of the center of gravity. Thus, all limitations on the use of modern wing mechanisms in the "duck" configuration are removed while maintaining static stability.
"KRASNOV-FLYGER"
Everything is great! However, there is a drawback. For positive lift to arise on the FGO 1, there must be negative lift acting on the servo rudder 3. This is analogous to a conventional aircraft design. In other words, there are losses in the balance, specifically the balance of the FGO. Thus, the way to eliminate this drawback is to use the "duck" scheme. We place the servo rudder in front of the FGO, as shown in Figure 3.
The FGO operates as follows (6). As a result of the aerodynamic forces acting on FGO 1 and servo rudder 4, FGO 1 spontaneously sets itself at a certain angle of attack to the direction of the oncoming flow. The angles of attack for FGO 1 and servo rudder 4 have the same sign, therefore, the lift forces on these surfaces will have the same direction. That is, the aerodynamic force on servo rudder 4 does not decrease but instead increases the lift of FGO 1. To increase the aircraft's angle of attack, the pilot shifts control rod 6 forward, causing servo rudder 4 to rotate clockwise around hinge 5, which increases its angle of attack. This results in an increased angle of attack for FGO 1, i.e., an increase in its lift.
In addition to controlling pitch, the connection provided by control rod 7 ensures an increase from zero to the required value of the FGO derivative.
Letâs assume the aircraft has entered an updraft and its angle of attack has increased. In this case, beam 2 rotates counterclockwise, and hinges 9 and 8, in the absence of control rod 7, would have to come closer together. Control rod 7 prevents this convergence and rotates servo rudder 4 clockwise, thereby increasing its angle of attack.
Thus, with a change in the direction of the oncoming flow, the angle of attack for servo rudder 4 changes, and FGO 1 spontaneously establishes itself at a different angle relative to the flow, creating a different lift. The magnitude of this derivative depends on the distance between hinges 8 and 3, as well as the distance between hinges 9 and 5.
The proposed FGO has been tested on the electric cord model of the "duck" scheme, with its derivative compared to the fixed GO being reduced by half. The load on the FGO was 68% of that for the wing. The goal of the test was not to achieve equal load distributions but to obtain a lower load on the FGO compared to the wing, as achieving equality would not be difficult. In "ducks" with a fixed GO, the load on the empennage typically exceeds the wing load by 20-30%.
The "Ideal Airplane"
If the sum of two numbers is a constant value, then the sum of their squares will be minimal when these numbers are equal. Since the inductive resistance of the lifting surface is proportional to the square of its lift coefficient, the minimum resistance of the airplane will occur when the coefficients of both lifting surfaces are equal during cruise flight. Such an airplane should be considered "ideal." The inventions of "Krasnov-duck" and "Krasnov-weathervane" allow the concept of an "ideal airplane" to be realized without resorting to artificial stability supported by automatic systems.
A comparison of the "ideal airplane" with a modern aircraft of a normal scheme shows that a 33% gain in commercial load can be achieved along with a 23% saving in fuel.
The FGO generates maximum lift at attack angles close to critical, and this mode is characteristic of the landing phase of flight. At this time, the airflow around the lifting surface is close to the boundary between normal and separated flow. The separation of flow from the GO surface is accompanied by a sharp loss of lift on it and, consequently, an intense nose drop of the airplane, known as a "nosedive." A notable case of a "nosedive" is the crash of the Tu-144 at Le Bourget, when it disintegrated during a dive after experiencing a nosedive. Using the proposed FGO easily solves this issue. It is enough to simply limit the servo control's turning angle relative to the FGO. In this case, the effective angle of attack of the FGO will be limited and will never equal the critical angle.
The "Weather Vane Stabilizer"
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The question of using the FG in a normal scheme is of interest. If instead of reducing, we increase the angle of the FG's rotation compared to the servo control, as shown in Fig. 4, then the derivative of the FG will be significantly higher compared to a fixed stabilizer (7).
This allows a significant shift in the focus and center of mass of the aircraft backward. As a result, the cruising load of the FG stabilizer becomes positive rather than negative. Furthermore, if the center of mass of the aircraft is shifted beyond the focus due to the deflection angle of the flap (the point where the lift increment occurs due to flap deflection), then the feather stabilizer also generates positive lift in the landing configuration.
However, all of this may hold true until we take into account the influence of braking and flow skew from the front lifting surface on the rear. It is clear that in the case of a 'duck' configuration, the role of this influence is significantly less. On the other hand, if the stabilizer 'carries' on military fighter jets, why would it stop 'carrying' on civilian aircraft?
'Krasnov plan' or 'pseudo-feather duck'
The articulated attachment of the destabilizer, although not drastically, does complicate the aircraft's structure. It turns out that the derivative of the destabilizer can be achieved through much cheaper means.
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Fig. 4 shows a stabilizer 1 rigidly linked to the fuselage (not shown in the drawing) of the proposed aircraft. It is equipped with a means to change its lift in the form of an elevator 2, which is attached to the bracket 4 via hinge 3, rigidly connected to the stabilizer 1. On the same bracket 4, a rod 6 is placed using hinge 5, with a servo control 7 rigidly fixed at the rear end. At the front end of the rod 6, near hinge 5, a lever 8 is rigidly fixed, the upper end of which is connected via hinge 9 to the rod 10. At the rear end of the rod 10, a hinge 11 connects it to lever 12 of the elevator trim 13 for elevator 2. The trim 13 is attached to the rear part of the elevator 2 via hinge 14. A coupling 15 changes the length of the rod 10 under the pilot's control for pitch management.
The presented destabilizer operates as follows. When the angle of attack of the aircraft randomly increases, for example, when entering an updraft, the servo rudder 7 deflects upwards, resulting in a shift of thrust 10 to the left, i.e., forward, which causes the trim tab 13 to deflect downwards, leading to the elevator 2 being deflected upwards. The positions of the elevator 2, servo rudder 7, and trim tab 13 in this scenario are depicted with dashed lines in the drawing.
As a result, the increase in lift of destabilizer 1 due to the increased angle of attack will be somewhat offset by the upward deflection of the elevator 2. The extent of this offset depends on the ratio of the deflections of the servo rudder 7 and the elevator 2. This ratio is determined by the lengths of levers 8 and 12. When the angle of attack decreases, the elevator 2 deflects downwards, and the lift of destabilizer 1 increases, offsetting the decrease in angle of attack.
Thus, a reduction of the destabilizer's derivative compared to the classical "duck" is achieved.
Since the servo rudder 7 and the trim tab 13 are kinematically linked, they balance each other. If this balance is insufficient, a balancing weight must be incorporated into the design, which should be placed either inside the servo rudder 7 or along the extension of rod 6 in front of hinge 5. The elevator 2 must also be balanced.
Since the derivative of the angle of attack of the lifting surface is approximately twice that of the angle of deflection of the flap, a twofold excess of the angle of deflection of the elevator 2 compared to the angle of deflection of the servo rudder 7 can lead to a value of the destabilizer's derivative close to zero.
The area of the servo rudder 7 is equal to that of trim tab 13 of the elevator 2. In other words, additions to the aircraft's design are quite small in size and negligibly complicate it.
Thus, it is entirely possible to achieve results similar to those of the "weather vane duck" using only traditional aircraft manufacturing technologies. Therefore, an aircraft with such a destabilizer can be called a "pseudo-weather vane duck." A patent has been obtained for this invention, titled "Krasnov-Plane" (8).
"Turbulence-ignoring aircraft"
It is quite reasonable to create an aircraft where the sum of the front and rear lifting surfaces has a derivative equal to zero.
Such an aircraft will practically completely ignore vertical air currents, and its passengers will not feel any 'bumpiness' even during intense atmospheric turbulence. Since vertical air currents do not lead to overload on the aircraft, it can be expected to have significantly lower operational overload, which will positively affect its structural mass. As the aircraft experiences no overload during flight, its airframe is not subject to fatigue wear.
The reduction of the derivative of the wing of such an aircraft is achieved in the same way as for the destabilizer in a 'pseudoflues' duck. However, the servo-rudder affects not the elevator, but the flaperons of the wing. A flaperon is part of the wing that functions as both an aileron and a flap. In this case, as a result of a random change in the angle of attack of the wing, an increase in its lift occurs at a focus based on the angle of attack. A negative increase in the lift of the wing resulting from the deflection of the flaperon by the servo-rudder occurs at a focus based on the angle of deflection of the flaperon. The distance between these foci is almost equal to one-fourth of the average aerodynamic chord of the wing. As a result of the actions of this pair of opposing forces, a destabilizing moment is formed, which must be compensated by the moment of the destabilizer. In this case, the destabilizer must have a small negative derivative, and the value of the wing's derivative should be slightly above zero. A patent for such an aircraft has been obtained (Russian Patent No. 2710955).
The combination of the inventions described represents, perhaps, the last unused informational aerodynamic resource for increasing the economic efficiency of subsonic aviation by a third or more.
Yuri Krasnov
REFERENCES
- D. Sobolev. A Century of the 'Flying Wing,' Moscow, Rusavia, 1988, p. 100.
- Y. Krasnov. Russian Patent No. 2000251.
- A. Yurkonenko. Alternative 'Duck.' Technology for Youth 2009-08. pp. 6-11
- V. Lapin. When Will the 'Flues Duck' Fly? General Aviation. 2011. No. 8. pp. 38-41.
- Y. Krasnov. Russian Patent No. 2609644.
- Y. Krasnov. Russian Patent No. 2651959.
- Y. Krasnov. Patent RF No. 2609620.
- Y. Krasnov. Patent RF No. 2666094.
Source: habr.com
