McDonnell-Douglas / Boeing F-15 Eagle

KJ_Lesnick

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The F-15 was able to fly on one wing because of it's wide-fuselage, and excellent control-power. I got a question, how much of a role did the variable-geometry inlet (you know the hinged inlet lip) lip position play in the amount of lift the fuselage itself produced?

KJ_Lesnick
Was the variable-geometry inlet lip design patented by McDonnell, or could any aircraft company use it?
 
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Lee

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KJ_Lesnick, quoted:
"Was the variable-geometry inlet lip design patented by McDonnell, or could any aircraft company use it?"

True, in my opinion. Was this something like what you were looking for?

http://www.google.com/patents?id=WOkHAAAAEBAJ&printsec=abstract&zoom=4&dq=variable+inlet+inassignee:McDonnell&as_drrb_ap=q&as_minm_ap=1&as_miny_ap=2008&as_maxm_ap=1&as_maxy_ap=2008&as_drrb_is=q&as_minm_is=1&as_miny_is=2008&as_maxm_is=1&as_maxy_is=2008&num=50

McDonnell Douglas patented that general design. Patents are meant to be wide in scope and American ones always include a disclaimer saying that many variations are possible in any patent.

However, anyone whosoever can claim improvements to a patent and then argue it out in court if sued over it. That should be how it works internationally.
 

KJ_Lesnick

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I thought the rule with patents was that if you modified it just a little you can get around the patent?
 
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Lee

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In theory, yes, minor variations in a patent should be allowable.

However, there should be either some improvement or useful benefit coming out of a patent compared to 'prior art'. That's the legal standard. Where the rub comes in is the opinion of what constitutes these virtues legally.

There's a whole field of legal specialization to Patent Law in America and all litigation is art---which really boils down to an opinion. Cultural art and legal arts are like that.
 

KJ_Lesnick

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I know the F-15 features a moving lip that can hinge up and down depending on speed which I'm guessing is to increase capture area at supersonic-speed. It also features a porous ramp.

My question is, is the porous-ramp fixed, or does it move as well (narrowing down the throat at high-speeds) in addition to the moving lip?


Kendra
 

rousseau

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Arh, my friend, if my understanding of which you mean porous-ramp is right, my answer is yes.
Only lip moving, will cause a turbolent flow in the inlet, so we need an incline to conform the air-flow after the lip down.
 

KJ_Lesnick

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rousseau said:
Arh, my friend, if my understanding of which you mean porous-ramp is right, my answer is yes.
Only lip moving, will cause a turbolent flow in the inlet, so we need an incline to conform the air-flow after the lip down.


I'm confused here exactly as to what you said.

It would seem you either said...
- 1.) Only the lip moves
- 2.) The lip moves with the ramp and the lip would produce turbulent flow in the inlet if only the ramp moves

Which one is correct?
 

rousseau

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Sorry, English is not my native language.
What I wanted to say is, both of them would moving responsively, not only one part moving factually. If you get a cutaway of F-15, you will find a board for release/discharge air pressure after the lip of inlet.

sorry for my clumsy express. :D
 

flateric

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That's how it looks
 

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KJ_Lesnick

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flateric,

Thank you for the drawing! It really helps when you can actually see it.


Kendra Lesnick
 

KJ_Lesnick

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Wait...

I just checked that patent again... those were dated 2001... did they patent the 'mission adaptive' idea (droopable inlet) when the F-15 was being designed and built? Or did the only do it in 2001?


Kendra Lesnick
 

Weaver

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Wasn't the first such intake of that general shape flown on the North American A-5 Vigilante? If so, then how do NA's patents (I presume they had some) compare with McDD's?
 

KJ_Lesnick

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I'm not talking about the ramps, I'm talking about the intake lip which droops at low-speed and opens up at high-speed.

Kendra Lesnick
 

KJ_Lesnick

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Out of curiosity, was the basic inlet concept, using the moveable-lids, patented by McDonnell Douglas around the time the F-15 was developed, built and flown?

Kendra Lesnick
(BTW: I know later on in 1999, they did patent the mission adaptive inlet, but I don't know if they patented anything before that)
 

flateric

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The thickness ratio varies from 5.9% at the
exposed root to 3% at the tip.


The F-15 wing development program
NIEDLING, L. G. (McDonnell Aircraft Co., St. Louis, Mo.)
AIAA-1980-3044
In: The evolution of aircraft wing design; Proceedings of the Symposium, Dayton, Ohio, March 18, 19, 1980. (A80-31001 12-05) New York American Institute of Aeronautics and Astronautics, Inc., 1980, p. 125-129.
 

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r16

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ı was about to write it was 6.6 at wingroot , thinning to 3 at the wingtip area . The source was a Klaus Huenecke book from 80s , but I guess the post by Flateric is the right one .
 

flateric

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I've seen number 6% at wingroot, and *exposed* wingroot is not the same, of course. Again, here we have original MacAir sourced paper...
 

flateric

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BTW, this is painful story of F-15 wing planform evolution
 

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weirc

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Hi this site http://www.ae.uiuc.edu/m-selig/ads/aircraft.html gives the root airfoil as NACA 64A 006.6 and the tip airfoil as NACA 64A 203 I have always found this to be a reliable site. With NACA 64 series airfoils the last 2 whole digits give the T/C ratio. On the drawing posted by Flateric the most inboard profle given is not the actual root airfoil but the one at the inboard end of the flap, so this could be the source of any confusion here.
 

KJ_Lesnick

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I take it the unexposed wingroot is the front of the wing-body fairing back to the tip of the booms on the back?

BTW: Just out of curiousity:
-What is that line down the airfoil (I think it looks like a conical camber thing or something, but I'm not sure)?
-The reason they deleted leading-edge devices from the design was because they were not necessary correct?


Kendra Lesnick
 

r16

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flateric , I am not doubting your source , it is just my typing . For all ı know we might be both right as ı understand wing areas are calculated to include parts what everybody would accept as the fuselage

what follows is an attempt to show how ı understand it and not an attempt to teach anything ; ı try to avoid to give such an impression even when I actually know anything about the subject ...

black lines are fuselage sides , wing area includes the yellow part inside the fuselage and my 6.6 would be at the blue line .

and the post by Weirc tells this might be the case . We are not disagreeing , just looking at the same thing from different perspectives .

about the book ; I don't understand at least %80 percent of it ...
 

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weirc

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Hi, I think everybody is correct on this one. The line labled BL 0.0 on Flateric's post is the center line of the aircraft, and, if the wing were continued to this point would be 6.6 percent thick. However the wing on the F-15 is only exposed to the airflow outboard of the inner end of the flap, so the most inboard part of the wing to be exposed to the airflow does have a thickness of 5.9 percent.

The line drawn along the airfoil in this case just shows the position on the wing of where the given airfoil section was taken from. If you were to see a straight line drawn from the leading edge radius to the tip of the trailing edge this would be the cord line, if you were to see a line drawn equally distant from the top surface and bottom surface of the airfoil that would be the camber line.

Hope it's getting clearer...........
 

KJ_Lesnick

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I have two more questions (I'm pretty sure I can't find an answer to these anywhere else and you guys know your stuff).

1.) What is the thickness of the F-15's wings at the thickest point and/or the unexposed root [/i](where it's at 6.6%)[/i], the exposed root, and around the area where the wing reaches 3% T/C?
2.) What is the chordwise length of the F-15's wings at the thickest point and/or the unexposed root (where it's at 6.6%), the exposed root and around the area where the wing reaches 3% T/C?


Also, if anybody knows these ones... (I'm not even sure you guys will know this one, but I might as well ask)

1.) What is the T/C ratio of the F-15U's (The F-15E derivative with the enlarged wings) wings?
2.) Does anybody know the chordwise thickness of the F-15U's wings at the exposed root, the unexposed root (if possible), and at the tips?
3.) Does anybody know the chordwise length of the F-15U's wings at the exposed root, at the unexposed root (if possible), and at the tips?


Kendra Lesnick
BTW: I have a good reason for why I'm asking all these extra questions... I'm actually working on a WHIF-design that is essentially an F-14/F-15 hybrid -- an "Eagle-ization of the F-14" if you will: The airplane basically is to look like a twin-seat F-15 style-design with the F-14's radome-size, the F-15U's larger wing area, the F-15A/C's wing-thickness, with droops and multiple-position flaps. Other ideas included a single all-moving vertical fin.
(I'm not trying to go off topic here... I just want people to know why I'm asking all these highly-detailed questions about the F-15 and it's wings!)
 

weirc

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Hi, if you want to know the thickness of an airfoil for a given cord and you have the t/c ratio of the airfoil, it's easy to work out. Measure the cord length of were ever you want on the wing, then, divide the answer by 100, then multiply that answer by the t/c ratio and that will give the size of the thickest point on the airfoil.

Sizes on airfoils, thickness, camber etc are often expressed as %'s of the cord so are usually straitforward to work out once you have the cord length for modeling purposes measuring form a decent scale drawing, to get the cord length, should be fine.
 

KJ_Lesnick

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weirc,

Not exactly as easy as you'd think. I can't find any really detailed 3-view drawings of the F-15U...


Kendra
 

KJ_Lesnick

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According to the book "F-15 Eagle Engaged" by Steve Davies and Doug Dildy, it states on page 19 regarding the September 1968 RFP requirements for the F-X that the fighter was among other things to be capable of achieving "Global (intercontinental) ferry range with or without aerial refuelling"

The exact passage reads as follows (It's a caption on the right side of the page)

F-X Design Requirements

The September 1968 RFP required the F-X design submissions to
provide a fighter with:


1. Wing optimized for high load factor (g) and buffet-free
performance at Mach 0.9 at 30,000ft altitude;
2. High thrust-to-weight ratio to achieve very high energy
maneuverability throughout the flight envelope;
3. Mach 2.5 maximum speed at altitude;
4. Long-range pulse-Doppler radar with look-down capability;
5. One man operation of the weapons system for all missions;
6. Advanced cockpit layout, displays, and controls, which would
allow heads-up operation during close-in combat;
7. Airframe fatigue spectrum with a life of 4,000 hours;
8. 360 degree cockpit visibility;
9. High maintainability: 11.3 maintenance man hours per flight hour
(similar to WW2 fighter requirements);
10. Significant increase in avionic and airframe subsystem component
mean time between failure (MBTF);
11. Highly survivable structure, fuel, hydraulic, flight control and
electrical subsystems in a combat environment;
12. Self-contained engine starting without need for groupd support
equipment;
13. Global (intercontinental) ferry range with or without aerial
refuelling;

14. Maximum air superiority mission gross weight in the
40,000lb class;
15. Low development risk components (engine and radar) and
airframe subsystems which had been proven in prototype,
pre-production, or production applications.



Source: Steven, James Perry McDonnell Douglas F-15 Eagle, Aero
Publishers, Fallbrook, CA, 1978.

Bold Emphasis Mine


How long, distance-wise, would an airplane need to be able to fly for the US Military to consider it intercontinental or global? Could anybody venture a guess as to whether this range would be with drop-tanks or on just internal fuel?

Was the plane able to actually meet this specification?


KJ Lesnick
 

red admiral

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I was having a look at the F-15 recently and found something that rather puzzles me. Hopefully someone with more experience of the aircraft can help with some explanation.

From the Standard Aircraft Characteristic, the Flight Design Gross Weight is listed as 37400lb at which 9g is the service g load. Above this weight, the service g load should reduce (37400*9/weight). So, at about 50,000lb, the service g load should be 6.7g

The chart below shows the Overload Warning System limits which are a function of Mach and Weight. From what I've been able to discern, this gives the pilot an audio tone that he is about to exceed the service g load. However, if we use the chart, say for Mach 0.85/sea level/50,000lb this gives 8.25g as the g load achievable before the warning system starts to tone. This doesn't agree with the above, as you're able to overstress the airframe by 23% before the warning kicks in.

Can anyone shed some light on this please?

Thanks very much
 

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kcran567

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The f-15 can pull up to 14 g's, and a few pilots have done that in combat without damaging the airplane, but then again I know of one F-15 that ripped its nose off at the fuselage connection probably due to metal fatigue from continued over stressing.
 

aim9xray

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A few thoughts to consider:

1: The chart that was posted is for a symmetric (non-rolling, non-yawing) maneuver.

2: The lower limit seen for the "rule-of-thumb" calculation may have sufficient margin to allow for asymmetrical maneuvers which cause different load patterns and concentrations in the airframe than the symmetric load. External stores also reduce the allowable margin.

3: The concern with an "over-G" is not just with damage for the primary airframe structure (i.e. wings "breaking off")- secondary elements such as landing gear uplocks, door hinges and engine mounts are particularly critical.

One of the YF-105As lost a main landing gear (uplock failed, the gear extended and was ripped off by airloads) while in a windup turn. The airframe was written off after a subsequent belly landing on the Edwards lakebed that broke it's back.

For a long time, the early Navy F-4Bs had a 6.5 G (symmetric) limit due to engine mounts. (The mounts themselves could not be strengthened alone; that would have just reacted the loads deeper into the primary structure which have been much more difficult and expensive to repair.)
 

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Bit off topic perhaps, but have Japanese and Israeli F-15's the same G-limit as USAF's F-15's? (C-version)
The same aircraft can have different G-limits with regard to the country it is operated by and requirements imposed by it's airforce. For example:
US' F/A-18C/D's have a 7.5 G-limit, while the Finnish' or Swiss' F-18C's (don't remember by heart if it are the Finnish or the Swiss, I'd have to look it up) have a 9 G-limit - with relatively minor reinforcements to the airframe. Their F-18's (not designated F/A-18's) are purely intended for A2A-combat and therefore have a higher manoeuver-requirement.
 

red admiral

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Thanks very much for the replies. You've given me some ideas to look into.
 

Mark Nankivil

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The initial Israeli Eagles were very early A models and in a couple of cases, reworked FSD Eagles so it may be possible they had different G limits for those airframes. As for the Eagle that lost its nose/cockpit section, that was a 131st FW Missouri Air National Guard aircraft from where I live (St. Louis). It was determined that a number of aircraft were built with longerons through the cockpit section (along the canopy rails) that were of a smaller cross section then were spec'd which led to fatigue cracking that in turn led to the loss of this aircraft. The fleet grounding found a few more with the same problem and in the case of one other with the squadron here, never flew again except for a one way trip to the boneyard.

The inflight break up is detailed here: http://www.youtube.com/watch?v=U22_7jsQy7s

Enjoy the Day! Mark
 

GTX

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Any ideas re this latest Israeli F-15 mod - specifically the hump behind the cockpit?
 

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