Why the small wing?

[sprstdlyscottsmn]

The F-16 has encountered phenominal growth for a fighter from the YF-16 stage untill the F-16E/F. This increase in structural weight, and avionics weight, and fuel weight, is to allow for more ordnance to be carried greater distances. The Engines are also upgraded to allow these heavier falcons to get off the ground. So why does the Falcon still carry the small 300 sqft wing? Why not up grade to a 350 sqft wing or somthing like what the F-2 has? I spoke to an AIr FOrce pilot who said a larger wing was planned for the Blk50 but was not implemented due to cost. input anyone?

[Roscoe]

1) A bigger wing would completely change the aero characteristics of the airplane requiring a huge flight test program
2) A 9g airplane puts tremendous loads on the wings. Bigger wings may spread out the load, but the bending moment at the root can actually increase.
3) F-2 had tremendous structural issues as a result of the bigger wing. USAF anticipated that and for that reason elected not to go with it. We recommended that Japan not...but they chose to listen to GD (LM?) instead.
4) More wing equals more drag.

[johnwill]

I agree that an F-2 wing is not the way to go, but not because it is larger. As member of the F-2 program, I can tell you the issues with the wing have little or nothing to do with the larger wing area. The customer wanted a "new" technology wing (among other things) to help justify the cost of the program. They acquired or developed new capability in flight control, avionics, structural materials, etc. The wing design used a new composite technology called co-curing, where the understructure (spars, ribs) and lower skin were bonded together into one component. All of the wing issues are related to the new technology, not the increased area. If the wing area had not been increased, the same issues would still have occurred.

Surface area increases are well within the capability of competent engineers. As Roscoe correctly states, cost and drag are the only disadvantages. Even drag is questionable and could be lower due to reduced AoA required. I offer several examples - the F-16 Block 15 horizontal tail and the F-16XL. Both had serious increases in area and both were totally successful structures. There are other examples - F-111B, F-111C and FB-111A from the basic F-111A, plus the F-16A and B wing and horizontal tail derived from the YF-16.

As far as increased bending moment goes, the F-16C/D Block 60 bending moment is much more than the F-16A/B, but has the same area. So there is no consistent correlation between area and moment.

[sprstdlyscottsmn]

more between pure weight and span for the moment?

[johnwill]

Yes. For example, flexibility. One of the differences between the F-16 wing and the F-2 wing other than the material (aluminum vs. composite) is flexibility. When air pressure load is applied to the wing it will both bend and twist. The twist changes the local angle of attack. A more flexible wing will twist more, so that the outer wing will be at a lower angle of attack than the inner (for a swept wing). The tip area will generate less lift, thus moving the center of pressure inboard, which means less bending moment. Although carbon fiber composite is stiffer than aluminum, it does not mean a carbon fiber wing will flex less than an aluminum wing. Carbon fiber is also stronger, so the skins do not have to be as thick as for aluminum. Which means the carbon fiber wing may in fact be more flexible than an aluminum one.

Associated with flexibility is the effect of sweep angle. More sweep means more negative twist, thus less bending moment. Forward swept wings will tend to have positive twist, thus more bending moment. The big hazard with positive twist is divergence, where more load means more twist which means more load and so on until the wing fails.

Mach number has an important effect on bending moment. The center of pressure on a wing shifts aft as mach increases (causing more negative twist) and reduced bending moment. At higher angle of attack, the inboard part of the wing will generally stall first, causing the center of pressure to move outboard - more bending moment.

Another difference is planform taper. A rectangular wing will have a higher bending moment that a tapered wing. External stores have a big influence on wing bending moments, as does the presence of internal wing fuel.

As you can see, it is not a simple matter to determine the strength requirements for a wing or any other part of the airplane.

[sprstdlyscottsmn]

Yeah I know what can go into it. I had to work on the lift and drag moments and sweep induced twist for a heavy lift cargo plane my senior year of school. We had an all composite wing but had a problem getting the skin onto the spars/ribs. I dont think that the sweep on the F-16 causes too much of a twist due to load when compared to a wing like on the F-86 of EE Lightning. Putting fuel in the outer sections of the expanded wing would help with the moments as would tip AMRAAMs.

I know increasing the wing area will increase the parasite drag, but the plan is to get it to reduce the induced drag enough to compensate. This of course would involve a new controll for the FBW.

Oh! If the F-2 wing was all composite did it offer any weight savings over the smaller aluminum wing? and with a ~20% increase in wing area, what kind of lift increase could be expected? Even if lift was created purely by the wing, which its not, the thickness would have to be the same for a 20% lift increase, so with a thinner wing (I assume F-2 had thinner wing ratio) and fuselage effects would it be closer to 10%? Sorry so many questions, I havent designed a fighter yet, only cargo.

Was the problem with the F-2 structure that it was composite, single piece, or trying to mate it to aluminum fuselage? Was the Fuselage even aluminum? Thank you johnwill for your input. I love the F-16 more than all other fighters and I always want to know conceptually how to make it even better.

[habu2]

Why? Money.

[johnwill]

Glad you are interested enough to ask good questions. I think you are right about twist effects compared to the F-86 and EE Lightning, but they still have to be considered. To show how powerful twist effects can be, the F-16 wing tip area is down loaded at low altitude supersonic level flight conditions. When you mentioned the problem of getting the skin load into the spars/ribs, you came close to one of the F-2 problems. As you know, the load between the skin and spars/ribs is shear. The lower skin is bonded to the spars/ribs, and that joint provided good shear strength. But at certain locations like the flaperon hinge ribs, the concentrated hinge load tries to pry the rib away from the skin, a tension load. Bonded joints are not very good in tension, so problems were found during ground test and redesign was required. Fuel in the wing doesn't help much, since it is the first internal fuel burned. That helps with roll maneuverability.

I am not familiar with weight comparisons, but wing weight is so small compared to maximum takeoff weight, it would not be very significant. The original F-16 wing structural weight was about 700 lb each, with a max weight of over 35000 lb.

Mating to the aluminum fuselage was no particular problem. Similar to the F-16, there are upper and lower aluminum brackets bolted to the wing and fuselage. There are also shear ties on the front and rear spars attached to fuselage bulkheads.

As to your original question, GD proposed an enlarged wing in the mid-80s to cope with current and expected weight increases. The AF had no money for that, so the wing has been redesigned for strength increase as required for block weight increases. Instantaneous g capability can be maintained that way, but not sustained g.

[sprstdlyscottsmn]

during our lift load test of our composite wing, human error reared its ugly head. We bonded the bottom skin to the ribs/spars with composite fillets but had to use a bizzar epoxy with composite chunks to bond the top skin. At about 1700lbs load the top skin buckled off the structure causing rapid failure of the system. lesson learned? Dont be too fancy by putting camber where it wasnt needed (the reason we couldnt test it upside down) and PAY ATTENTION to which side gets the stronger bond. Had we filleted the upper skin the buckling resistance would have been phenominal. The poor bond on the lower skin would have bad little effect as that side was in tension.

Back to the subject. I can see how flaperons would complicate the issue. Our test was just of primary structure so ailerons were not modeled (mostly for time/money) but I could see shear posing an issue there. I did not know that the tips were down laoded during low level high speed flight. How much natural twist does the Vipers wing have?

Did putting holes for bolts in a compostie structure create any issues for delamination? I have only a rudimentary knowlege of composites.

Another topic, a pure delta at high alpha (say M-2000 at 30 deg) the upward deflection of the elevons creates the pitch up moment to gain alpha right? So if a roll input (left stick) is performed would the increase of lift on the wing with the lowered elevon(right wing) over come the reduced pitch moment. What effect does a differential pitch moment have on a plane in that situation?

[johnwill]

The test you describe is very interesting. Upper skin buckling is always a problem. That is why the upper skin is normally thicker than the lower. The F-16C static test wing failed (like you described) at 137% of limit load due to buckling (150% is required). Since composites are stronger than aluminum, the composite skin can be thinner to resist bending moment, but then the buckling problem becomes much worse. Composite upper skins many times have more spars or other stiffening devices for stiffening the skin to resist buckling. When you talk about the lower skin bond having little effect since it was in tension, only the lower skin is in tension - the bond is primarily shear.

The static twist on the F-16 wing is -3 degrees. So if you are flying at 4 deg AoA with no elastic twist, the wing tip still has positive AoA. But as you approach and pass 0.9 mach, two things happen - AoA goes down due to higher dynamic pressure (q) and the center of pressure moves aft. If AoA is 3 deg and no elastic twist, tip load is zero, but the aft cp means there is elastic twist, making the tip AoA negative. Once that happens, the tip missile airload is also down. Since it sticks out far ahead of the elastic axis, that twists the wing even more nose down. It is very dramatic for the pilot to look at his tip missiles and see they are pointing several degrees nose down. The lift distribution on the wing is strange - down load on the tip, gradually going to zero then positive as you move inboard. The total lift (shear) is positive, but the root bending moment is negative.

One of the reasons for co-curing the F-2 wing was to avoid holes in the lower skin. I also am no expert on composite structure, but is seems reasonable that delaminations are more likely to start at a hole or any edge rather than away from an edge. Maybe not. Another reason is to reduce fuel leaks.

Rolling a delta wing airplane could cause some problems with pitch unbalance if the pitch elevon deflection is near maximum. With small pitch elevon deflection, roll commands could be sent equally to left and right elevons, thus maintaining pitch deflection. If the pitch deflection is large, a roll command could cause one elevon to reach its travel limit while the other elevon continued moving. Pitch unbalance would result, and the nose would drop.

Sample numbers - de max = +/-30 deg

Before roll de (L) = -10 de (R) = -10 Pitch de = -10
Roll command +/- 15 de (L) = 5 de (R) = -25 Pitch de = -10

But if
Before Roll de (L) = -20 de (R) = -20 Pitch de = -20
Roll command +/- 15 de (L) = -5 de (R) = -30 Pitch de = -17.5

Since de cannot exceed +/- 30 deg, the right elevon is travel limited, and de is reduced (the nose goes down)

This case is similar to the F-16 rolling with gear and flaps down, where only one flaperon moves in a roll command, reducing total lift. The test pilots said it felt like the airplane was rolling around an outboard x-axis. They soon became accustomed to it and it was never an issue.

Also in the F-16, horizontal tails are used to help roll the airplane. Below about 0.9 mach they don't help much, but as the flaperons become less effective at higher mach, the tail roll commands are increased and the tails provide most of the roll power. The same condition you describe for the Mirage 2000 could possibly occur in the F-16 tails. In this case the filght control people have given priority to pitch control over roll control. If the roll commands could cause limit tail movement, the roll commands are reduced so that pitch command is never limited.

[sprstdlyscottsmn]

"The same condition you describe for the Mirage 2000 could possibly occur in the F-16 tails. In this case the filght control people have given priority to pitch control over roll control. If the roll commands could cause limit tail movement, the roll commands are reduced so that pitch command is never limited."

hence why rolling g is significantly lower than symetric G?

"The lift distribution on the wing is strange - down load on the tip, gradually going to zero then positive as you move inboard. The total lift (shear) is positive, but the root bending moment is negative. "

Man I want to see a chart of that, sounds crazy but I can see it. That moment arm out at the tip simply overpowers the net positive shear force. would that be one of those times that the body produces significant percentage of lift?

[johnwill]

The conditions involving possible full deflection of control surface occur at very low airspeeds, where the airplane can't pull more than 2g anyway. F-16 roll g limit (6g clean) is lower than symmetric g limit (9 clean) primarily due to historical precedent. With strictly manual control systems, rolling at 1g was easily controlled, but as g increased control was more difficult and roll performance was degraded. At high AoA, roll commands turn into yaw commands and the airplane may exceed its yaw stability limits.

So the standard way to perform high rate rolls was to reduce g. With the arrival of jet engines and ever higher speeds, the problems became worse. Even as late as the F-4, roll performance was much better at lower g. So the AF normally specifies max roll g as 80% of max symmetric g. Obviously, the F-16 is different, because 6g is less than 80% of 9g. All US fighters up through F-16 were specified to be 7.33g symmetric and 5.86g roll. (I am not sure about F-22 and F-35 roll g, but I think it may be 7.2g). During development of the F-16, GD proposed adding 9g capability at a very low weight penalty (22 lb), so the AF bought it. However max roll g was left at 5.86g. High g rolls are difficult conditions for the structure to withstand and would have required extensive redesign. After flight test confirmation, the limit was raised to 6g primarily because it is easier to observe. Todays control systems make it easier to control yaw with automatic aileron/rudder interconnects. F-16 roll rates at 6g are not much different than at 1g.

Wing structure is primarily designed by bending moment and torsion (twist). Symmetric maneuvers are mostly bending moment and 1g rolls are mostly torsion due to flaperon load. 6g rolls are a combination of bending and torsion as you would expect. Of course in the roll one wing will have added bending and the other wing will have reduced bending. How much? Depends mostly on mach number and its effect on flaperon efffectiveness.

Well, I've gone on far too long on this topic and have only scratched the surface.

You've got it exactly right about wing load distribution with negative tip load. Fuselage lift is always significant, but it becomes more of a contributor at lower speeds. Look at it this way - the wing lift vs AoA is curved, so that it is reduced significantly above 12 - 14 degrees. Yet for some reason, the fuselage lift curve is almost linear up to about 20 deg AoA. So for high g conditions with AoA < 12 degrees, the fuselage lift is around 40% of the total. But at say 18 deg AoA, the wing has lost some of its capability and the fuselage is still doing its job at perhaps 50% of total lift.

Thanks for asking good questions.

[sprstdlyscottsmn]

Thanks for giving good answers.

[sweetpete]

Hey, your guys brains seem bigger than mine so ill ask a question not a single F-16 pilot i have asked has been able to tell me. What is the primary purpose of the Anhedral in the stabs? From my limited understanding of fixed wing aerodynamics dihedral in a wing is designed to help an aircraft return to a neutral "level" plane, as the wing lowered during the maneuver becomes closer to horizontal its lift vector becomes more vertical than the opposing wing thus righting the aircraft by design. Based on this theory Anhedral in the stabs would serve to make the aircraft more unstable in the longitudinal plane making the maneuver more pronounced (better roll rate)? Sorry for going off topic but I can see the one of you posting on this topic prolly has the answer.

[johnwill]

The horizontal tails droop 10 degrees to provide more directional (yaw) stability. Look at the F-4 for an extreme case of the same idea. You can think of the tails as having a horizontal component (tail area x cos 10 deg) and a vertical component (tail area x sin 10 deg). The vertical component acts just like a vertical tail or a ventral fin - helps to keep the pointy end of the airplane in front. The vertical tail can be smaller (less drag, weight, cost) by doing this. Very little is lost in the horizontal component (cos 10 = .985), while a good gain is made in the vertical component (sin 10 = .174).