Patent Grant
10850827
Not Applicable.
The purpose is to improve the airfoils' performance for the construction of wings both in their low-speed performance (takeoff and landing) and their stability at high speeds with an acceptable capacity for inverted flight (acrobatic capacity).
It is well established in the field of aviation and aeronautics that airfoils are developed in accordance with specific purposes of flight (of speed, for gliding, acrobatic, etc.). The objective when designing these airfoils is to have airfoils with an improved flight capacity both at low and at high speeds, and with the capacity for inverted flight. There might be in the state of the technique documents such as the U.S. Pat. No. 6,607,164 B2, which presents an airfoil having particular use in a general aviation aircraft operating at generally low speeds. Said airfoil has a shape designed to produce high lift coefficients at low speeds and low drag and lower lift coefficients at higher speeds. Said airfoil's characteristics are not sensitive to surface roughness on the leading edge caused by the accumulation of foreign matter on the airfoil due to the transition to turbulent flow occurring near the leading edge at high lift coefficients, therefore limiting effective use of the airfoils disclosed in the '164 patent to low speeds. In contrast, the airfoils and wings disclosed herein are intended to be used at low speed, high speed and have capacity for inverted (aerobatic) flight.
Airfoils jn1432-265 and jn1413-362, as shown in
When analyzing the scale effect we found that the different curves when going through angles between 0 and +1 (taking into account that the angle of incidence in which the wing normally flies is within this range) of the graphs, the lift coefficient is highest when the Reynolds number is lowest and decreases as the Reynolds number increases. As the Reynolds number increases the coefficient adjusts to each flight condition, therefore the coefficient is high at slow speeds allowing for short, predictable and safer takeoffs and landings. The coefficient decreases as the speed increases which creates stability by allowing for more flexibility in different flight conditions. It has also been observed that UAVs configured with these wings for testing have a better performance in conditions with increased winds as compared to aircraft that have been configured with other airfoils. Also, the drag coefficient (cd) which in itself is low in the highest values of the lift coefficient (cl) also descends to values up to one third of the initial value as the Reynolds number increases.
The following table contains the coordinates of airfoil jn1431-265 which will be used for the wing root because it has the lowest lift coefficient and allows for the most stable stall.
The following table contains the coordinates of airfoil JN1413-362 which will be used for the wing end.
Airfoil jn1431-265 as the wing root combined with airfoil jn1413-362 as the wing end create the aforementioned characteristics of wing performance.
Airfoil jn1431-265 is 14.31% wide in relation to its length and airfoil jn1413-362 is 14.13% wide in relation to its length. Airfoil jn1431-265 has a camber of 2.65 and airfoil jn1413-362 has a camber of 3.62. Airfoils jn1431-265 and 1413-362 operate intelligently by adjusting their variable aerodynamics, not only by the angle of attack, but also by the scale effect (speed), as shown in
Consider the airflow on the airfoil of an airplane wing. With a determined angle of attack the airspeed is supposed to be uniform. The air in free flow (far from the airfoil) will be considered with properties as air at sea level: pressure=101.325 Pa, density=1.2250 kg/m3, temperature=288.16K, kinematic viscosity=1.4607E-05 kg/ms.
Under these conditions, the lift and drag coefficients were determined by using FLUENT for airfoils jn1431-265 and jn1413-362, and for comparative purposes airfoil Selig 1223 (S1223) was added; by function of the angle of attack and the Reynolds number. Additionally the net lift force was obtained and the maximum load which an established surface area can support was determined.
A wing was used measuring 1.524 meters long (L), with a chord (c) measuring 0.3048 meters. The weight of the wing (w) was 3.587 kg, the weight of the fuselage (wf) was 5.702 kg. The minimum load (Im) was 0.861 kg. And the wing's surface area was: A=1.3328 m2.
A simulation domain was set to be sufficiently large so that the far-field boundaries are sufficiently far away from the object causing the flow disturbance, in this case the airplane, and this way the result will be more exact.
The domain defined by the far-field boundaries is shown in
Referring to
The following table depicts the boundary conditions for the simulation:
The following chart depicts the velocity conditions of boundaries farfiled1 and farfield2 for the simulation:
The Reynolds number, Re, was calculated by using the following formula:
ρ represents density, μ viscosity and Uo represents the magnitude of free flow speed and c is the chord length of the wing's airfoil. X-vel and y-vel are the components of free flow speed in x and y. AoA (a) corresponds to the angle of attack.
The following chart depicts the draft and lift from the domain simulation for each angle of attack:
It was considered that the flow was isothermal given its low speed. On the other hand the turbulence was considered in order to obtain more exact lift and drag coefficients. For this the Spalart-Allmaras model was chosen, which was specifically designed for aerospace applications and is suitable for flows with boundary layers subjected to adverse pressure gradients. The PRESTO algorithm was used for the pressure equation and the Simple algorithm for coupling the speed and pressure equations. All the calculations were made on Double Precision. The convergence was monitored through residuals normalized for each equation. A 1E-4 convergence criterion was utilized for pressure and 1E-6 for the other variables. A second-order upwind discretization scheme was used for the momentum and turbulence equations.
The following chart depicts the characteristic values for calculating lift and draft coefficients:
Referring to
Referring to
When comparing the results presented we observe that the model with the highest lift coefficient is airfoil jn1413-362 followed by airfoil jn1431-265 and lastly Selig S1223. The latter for comparison purposes.
In order to calculate the maximum load weight the lift force (FL) will be calculated, which formula is:
FL=clρUo2A
A is the wing's surface area. The tridimensional lift coefficient is Cl max=0.9cl.
The maximum load weight is obtained with the formula:
Where g is the acceleration of gravity. And the weight of the airplane is obtained by adding the weight of the fuselage, of the wing and the minimum load. The results are shown on table of
The following chart depicts the maximum load results:
Airfoils jn1413-362 and jn1431-265 can carry substantially more weight than airfoil S1223 (which was used for comparative purposes). Even though each airfoil may be used separately in building a wing; the combination of airfoils jn1413-362 (for the wing tip) and jn1431-365 (for the wing root) given their differential in the lift coefficient contributes to a better lateral stability.
| Number | Date | Country | Kind |
|---|---|---|---|
| MX/u/2014/000421 | Aug 2014 | MX | national |
This application is a continuation in part of U.S. patent application Ser. No. 15/504,625.
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| Number | Date | Country | |
|---|---|---|---|
| 20200070950 A1 | Mar 2020 | US |
| Number | Date | Country | |
|---|---|---|---|
| Parent | 15504625 | US | |
| Child | 16571333 | US |
