Laminar flow wing optimized for transonic cruise aircraft
View Patent ↗Aircraft configured to operate at Mach numbers from above 0.80 and up to 1.2 with wing sweep angles defined by the wing outboard leading edge of less than 35 degrees, and incorporating calculated values of the ratio of outboard wing panel aspect ratio raised to an exponent of 0.78, divided by the ratio of maximum thickness divided by chord (t/c), greater than about 45, characterized by one of the following: a) where maximum thickness divided by chord (t/c) is at a location approximately 70% of the distance outboard from the attaching aircraft body to the wing tip, or b) where maximum thickness divided by chord (t/c) is the average value of (t/c)'s located between approximately 50% of the distance outboard from the attaching aircraft body to the wing tip.
1. Aircraft configured to operate at Mach numbers from above 0.80 and up to about 1.2 with wing sweep angles defined by the wing outboard leading edge of less than 35 degrees, and incorporating calculated values of the ratio of outboard wing panel aspect ratio raised to an exponent of 0.78, divided by the ratio of maximum thickness divided by chord (t/c), greater than about 45, characterized by one of the following: a) where said maximum thickness divided by chord (t/c) is at a location approximately 70% of the distance outboard from the attaching aircraft body to the wing tip, or b) where said maximum thickness divided by chord (t/c) is the average value of (t/c)'s located between approximately 50% of the distance outboard from the attaching aircraft body to the wing tip.
2. The aircraft of claim 1 with wing shapes configured to provide pressure gradients favorable to generating extensive laminar flow on upper, lower, or both wing surfaces.
3. The aircraft of claim 1 with t/c ratios substantially greater at locations inboard of the 50% wing span location as defined in claim 1 than at other locations.
4. The aircraft of claim 1 , combined with fuselage body and/or engine nacelle bodies with cross-sectional areas reduced adjacent to the wing or wings so as to reduce Mach related wave drag, such areas being in planes normal to flight direction.
5. The aircraft of claim 2 where the wing is defined as a tail or forward canard stabilizing or control surface.
6. The aircraft of claim 3 where the wing is defined as a tail or forward canard stabilizing or control surface.
7. The aircraft of claim 4 where the wing is defined as a tail or forward canard stabilizing or control surface.
8. The aircraft of claim 1 wherein the t/c ratio is greater than 0.03.
9. The aircraft of claim 2 wherein the t/c ratio is greater than 0.03.
10. The aircraft of claim 3 wherein the t/c ratio is greater than 0.03.
11. The aircraft of claim 4 wherein the t/c ratio is greater than 0.03.
12. The aircraft of claim 5 wherein the t/c ratio is greater than 0.03.
13. The aircraft of claim 1 wherein the t/c ratio is less than 0.08.
14. The aircraft of claim 2 wherein the t/c ratio is less than 0.08.
15. The aircraft of claim 3 wherein the t/c ratio is less than 0.08.
16. The aircraft of claim 4 wherein the t/c ratio is less than 0.08.
17. The aircraft of claim 5 wherein the t/c ratio is less than 0.08.
18. The aircraft of claim 6 wherein the t/c ratio is less than 0.08.
19. The aircraft of claim 1 characterized by
i) maximum design cruise Mach number greater than 1.2, and
ii) panel Ap^0.78/(t/c) 70 exceeding 67.
20. The aircraft of claim 19 characterized by wing leading edge sweep angularity less than 35 degrees.