Method and device for decreasing the flow resistance on wings particularly aerofoils and blades of turbomachines exposed to gas flux such as air

A device and method for reducing the flow resistance of air foils utilizes one or more flow compressors on the leading edge of the foil. The gas flowing in front of the leading edge is compressed and is expanded on the upper surface of the wing so that through induction of the expanded gas, gas from the area of the lower surface of the wing is admixed to increase the flow on the upper surface of the wing. The direction of the admixed gases is such that the velocity of the gas flowing out at the trailing edge of the wing is identical over the entire wing length. The separation point of the flow is displaced towards the trailing edge of the wind and, in the direction of the tip of the wing, the pressure of the gas flowing over the upper surface of the wing is increased.

The present invention relates to a method for reducing the flow resistance 
of wings, particularly airfoils and blades of turbomachines around which 
flow gases, such as air or the like, and to a device for performing the 
method. 
When air or some other gaseous fluid flows around the wings of the 
aforementioned type, the problem exists that a flow can only be obtained 
with a limited resistance if in the vicinity of the wing profile vortex 
formations are avoided or at least minimized, because flow eddies 
considerably increase the flow resistance of the wing. In the case of 
airfoils the problem exists in that in the approach flow area of the 
profile, the laminar interface changes over from a laminar into a 
turbulent flow over the profile depth, so that in the rear area of the 
airfoil profile a vortex area is formed. In addition, at the edge of the 
airfoil pressure differences in the flow on the upper and lower surfaces 
of the airfoil lead to marginal vortexes, which slope downwards towards a 
vortex train. In order to limit the effects of these features, it has 
already been proposed to arrange special flaps on the leading and/or 
trailing edge of the wing and by means of which energy is supplied to the 
flow on the top of the profile by induction action. Although this energy 
supply displaces the separation point of the flow towards the trailing 
edge of the wing, it cannot eliminate it. A particular disadvantage of 
these flaps is that when they are extended the fineness ratio of the 
airfoil profile is in part considerably impaired. For preventing marginal 
vortexes it is also known to provide vertical auxiliary wings at the wing 
ends which, although reducing vortex formation, increase the flow 
resistance. In order to minimize the flow losses in the case of blades and 
which consists of the separation and turbulence losses, it is known to 
twist or warp the blades over their length, in order to prevent a 
breakaway of the flow from the blades and consequently turbulent vortex 
formation by an oblique incidence of the flow in the front area of the 
profile. In the rear area and at the free end portion of the blades, 
turbulence occurs in the same way as on the airfoil and the magnitude 
thereof is dependent inter alia on the flow velocity. It has therefore 
already been proposed to pivot the blades in the vicinity of the blade 
base in order to counteract breaking away of the flow by varying the angle 
of incidence. However, this requires high constructional, manufacturing 
and control expenditure and increases the failure property of a 
turbomachine equipped with such blades. 
The object of the invention is to provide a device making it possible, 
without using special flaps, to influence the flow round profiles of 
wings, particularly airfoils and blades of turbomachines in such way that 
the separation point of the laminar wing length is displaced towards the 
trailing edge of the wing. According to the invention this object is 
achieved in that gas flowing in front of the leading wing edge is 
compressed and is expanded on the upper surface of the wing, so that 
through induction of the expanded gas, gas from the area of the lower 
surface of the wing is admixed and, simultaneously, the flow on the upper 
surface of the wing is increased and directed in such a way that the 
velocity of the gas flowing out at the trailing edge of the wing is 
identical over the wing length. The separation point of the flow is 
displaced towards the trailing edge of the wing and, in the direction of 
the edge of the wing, the pressure of the gas flowing over the upper 
surface of the wing is increased. 
According to a further development of the invention, the device for 
obtaining the compression ratios sought on a wing according to the 
invention is characterized in that one or more blades are arranged on the 
top surface of the wing in the vicinity of its leading edge and are 
integrated into the profile of the root of the wing and are uniformly 
twisted over its length in such a way that 
(a) the distance between the blade and the wing and optionally between the 
blades is widened to a gap of maximum width at the edge of the wing, 
(b) the mean camber line of the blade has a maximum camber at the root of 
the wing and is flat at the free edge, 
(c) and the chord of each blade to the chord of the wing is generally 
parallel from the base of the blade to the edge thereof.

The element 1 has the general configuration of a blade and is twisted over 
its length (FIGS. 1 and 2). As shown in FIGS. 3 to 5, the blade 1 is 
arranged on the top surface 3 of the wing in the vicinity of its leading 
edge 4. Blade 1 has a negative angular sweep towards the leading edge 4 of 
the wing that the trailing edge 14 of the blade 1 is located in the 
vicinity of its free end portion 15 in the flow direction in front of the 
leading edge 4 of the wing. As a result of the uniform twisting of blade 1 
over its entire length the distance between blade 1 and wing 2 is widened 
to a maximum width gap 6 at edge 7 of wing 2. A root 5 of wing 2, the mean 
camber line 8, see FIG. 6b, of blade 1 has a maximum camber whereas it is 
flat at the free edge. Chord 10 of blade 1 is oriented in a negative angle 
to chord 11 of wing 2 at base 12 of blade 1 and in a positive angle at the 
edge of blade 1 (FIGS. 6 a to 6c). 
Base 12 of blade 1 is mounted in a recess 16 of wing 2, where it is fixed 
by means of screw connections 17. Recess 16 is constructed in such a way 
that it is possible to embed base 12 of blade 1 in a flow-favourable 
manner. It is also possible to pivotably adjustably mount blade 1 in 
recess 16 by means of an adjusting device. This makes it possible to adapt 
the position of blade 1 for different operating conditions. Furthermore it 
is possible to provide on wing 2 a plurality of blades 1 having a negative 
sweep in the approach flow direction, in order to achieve further thrust 
increase of wing 2. 
Through the use of the blade 1 on wing 2 is formed in each case one 
underpressure zone 18 and overpressure zone 19 advanced towards the 
leading edge 4 of the wing. The overpressure zone 19 is formed on the 
bottom of the blade 1 and leads to a pressure compensation through gap 6 
by means of the outflowing medium. This pressure compensating flow induces 
a further gas supply from pressure zone 20 below wing 2, which is admixed 
with the inflowing gas. The gas flowing on the top surface of with wing is 
so oriented by pressure compensation that the velocity of the gas flowing 
away at the trailing edge of the wing is the same over its entire length 
and the separation point of the flow is displaced towards the trailing 
edge of the wing (FIG. 7). In addition, the pressure of the gas flowing 
over the upper surface of the wing is increased in the direction of edge 
7. This leads to a reduction of the pressure gradient and edge 7 and 
consequently to the formation of smaller vortex trains 24, so that the 
induced resistance of the overall profile is reduced. In the vicinity of 
the edge, an area of increased flow velocity is formed, so that the 
separation point in the turbulent interface is moved out over the profile 
of wing 2. Simultaneously the vorticity field at the rear wing area is 
reduced. 
Through the construction of blade 1 and its arrangement on wing 2, it is 
consequently possible by energy supply to compress the gas flow located on 
the upper surface of the wing and to increase the area of laminar flow; 
whilst reducing the vorticity flows, the result being that the thrust 
obtainable is increased. 
It is also possible to arrange one or more blades according to the 
invention on a wing 2 which, in per se known manner, is rotatable by means 
of an adjusting device for varying the angle of incidence at the base or 
root of wing 2. As a result it is possible to still further increase the 
efficiency improvement of the profile of wing 2 obtained by adapting the 
angle of incidence to the ideal approach flow, because here again the 
blade or blades improve the flow round the profile of wing 2 by minimizing 
vortex formation. As a result of the individual adjustment possibility of 
each blade, the flow round the profile can be further optimized in the 
case of wings 2 variable with respect to the angle of incidence. 
FIG. 8 shows a wing 2, which is connected to two blades 1, 1a fixed in a 
recess 16a on root 5 of wing 2. They are arranged in such a way that they 
extend forwards with respect to the leading edge 4 of the wing. In 
addition, edges 9, 9a of blades 1, 1a are raised towards the top of the 
profile with respect to edge 7 of wing 2. As a result between the blades 
1, 1a and wing 2 free flow cross-sections 27 are formed in the approach 
flow direction and pressure compensating cross-sections 26 at right angles 
to the approach flow direction. As shown in FIGS. 9a to 9c, wing 2 and 
blades 1, 1a are also cambered. The camber has a maximum in the vicinity 
of root 5, whilst the cross-sections of blades 1, 1a and wing 2 are flat 
at edges 9, 9a, 7. In this area blades 1, 1a are almost symmetrical. 
Whereas the chords 10, 10a of blade 1, 1a are at a positive angle to chord 
11 of wing 2 at root 5, chords 10, 10a of blades 1, 1a are approximately 
parallel to the wing chord 11 in the vicinity of edges 9, 9a due to the 
prevented profile cambering. This arrangement of the blades 1, 1a on a 
wing 2 shows in FIGS. 8 and 9a to 9c permits a particularly great 
efficiency increase. 
It is also possible to use the aforementioned flow blades 1, 1a for 
increasing the efficiency of marine propellers. In this case a blade 1 or 
several blades 1, 1a is or are associated with the propeller blade. In 
general each blade 1, 1a can be curved in accordance with the leading edge 
of the associated propeller blade. Through the use of blades 1, 1a the 
risk of cavitation of such propellers is considerably reduced, because 
critical underpressures at the propeller edges are avoided. It is also 
possible to use blades 1, 1a are associated with each propeller blade. 
Through the use of blades 1, 1a the problems caused by local sound 
velocities in the edge area of propellers are avoided.