# Why are there no cargo aircraft with “flying wing” design?

From that I have seen so far, the "flying wing" design (like the one of B-2 Spirit and Northrop YB-49) has superior performance but also a few notable problems that make it difficult to use for passenger aircraft:

• It is difficult to control, and the YB-49 crashed even when flown by an elite test pilot. However, computer assistance has been implemented for B-2 and I do not think this is a problem any longer.
• There are problems related just to the passenger transport: not enough windows, difficult to evacuate.
• It also cannot be pressurized as easily as a cylinder but for a majority of possible cargo this is probably not a problem. Some cargo may not require pressurization at all and some may only need partial pressurization like in jet fighters.

Hence I understand that there are problems on the way to the flying wing passenger aircraft. However, why there are no cargo aircraft of this kind around?

• – Peter Kämpf Apr 15 '19 at 13:36
• "computer assistance has been implemented for B-2 and I do not think this is a problem any longer" Boeings have been computer assisted for years (decades) and even they still have problems. An airplane being computer assisted isn't the end of all problems and it's not a magic bullet. – Mast Apr 18 '19 at 0:00
• "It also cannot be pressurized as easily as a cylinder but for a majority of possible cargo this is probably not a problem" True, but since you mention passengers a couple of times, it's quite a problem for passenger transports at those operational heights and speed. – Mast Apr 18 '19 at 0:01

## 10 Answers

Flying wings can be made to have acceptable flying qualities without any artificial assistance. Just look at the Jim Marske glider designs.

The principal downfall of flying wings is that stability in pitch is pretty much achieved the same way as with a conventional tail, with a down force balancing out the center of gravity forward of the fulcrum of neutral point of the lifting forces, but it's all being done over the very short moment arm of the wing chord itself. In other words the "tail" has been moved forward to the trailing edge of the main wing.

There are a lot of issues that result from this, pitch sensitivity and damping issues and all that, but the biggest one from a cargo aircraft's perspective is a very narrow center of gravity range. Not a big deal on a bomber with a concentrated bomb bay load, or a glider that doesn't have to cope with loading variations, but a bigger deal on a freighter. You are forced to spread the load, and the fuselage volume, laterally, creating way more frontal area than necessary (you're in effect turning the fuselage sideways), so you end up cancelling out the drag benefit of doing away with the tail in the first place, and still end up with a "temperamental" configuration.

• Admittedly, without a long fuselage there will not be much length along which the cargo can be distributed. I'd call it a wash. – Peter Kämpf Apr 15 '19 at 18:18
• That's what I meant by having to spread the loading laterally. But even within the space envelope you would have just within a flying wing stump fuselage or center section, the available loading range is pretty narrow. Bring your knees to your chest in a FW glider, where the allowable range is couple of inches, and you might find yourself aft of the rear limit. – John K Apr 15 '19 at 21:34
• most excellent explanation! – niels nielsen Apr 16 '19 at 1:50
• This explanation is plainly wrong, as stability is not due to this. an aircraft can be perfectly statically stable with the center of lift ahead of the center of gravity. - In fact many aircraft work that way and it's more stable that way. - This is due to the lifting moment, and the way cl-alpha works. – paul23 Apr 16 '19 at 23:28
• @paul23 the "flying wing" will still have a narrower CG range. And the point about "turning the fuselage sideways" is also correct. But good to hear from someone about the weight ahead of CG/down force on tail design. This ties in well with the concept of making tails smaller in a safe manner, rather than stalling tiny Hstabs in down wash because weight is unnecessarily too far forward. – Robert DiGiovanni Apr 17 '19 at 0:47

Cargo aircraft (outside the military) almost always started life as passenger aircraft. The ratio of active large cargo aircraft to passenger aircraft is in the single percentages. Therefore, nobody develops a pure cargo aircraft from scratch.

That does not mean that no one has tried. Especially for cargo, large flying wings have been proposed which store their cargo in containers along the wingspan - hence their name: Spanloaders. Below is an artist impression from the 1970s.

Boeing Model 759-159 distributed load freighter concept from the 1970s (picture source)

• And to the military, soldiers are just another kind of cargo. – jamesqf Apr 15 '19 at 17:03
• Where would that thing park? – Azor Ahai -him- Apr 15 '19 at 18:30
• Not at the airport it's flying over, certainly... – Roger Lipscombe Apr 15 '19 at 18:53
• Perhaps it doesn't park, or even land - just flies endlessly while smaller craft ferry fuel and cargo between it and the ground. – Skyler Apr 15 '19 at 20:08
• @RogerLipscombe; Actually, if you look closely, you can see two of them on the ground. But I agree, a standard 98 ft runway will be too narrow. – Peter Kämpf Apr 16 '19 at 7:07

For a start, with what it costs to design and certificate a new aircraft type, if a transport craft can't be reconfigured to carry either passengers or freight it won't make it off the napkin. The conventional transports we have can be switched from cargo to passenger and back, some in just a few hours. For a non-passenger transport to compete, it would have to be much cheaper (to buy and to operate) than a multi-purpose airframe.

In addition to the other answers, a reason for the lack of flying wings in civil aviation in general is that they need to compete in an environment that has grown alongside conventional, fuselage-and-wings aircraft and is ill-suited for flying wings.

This means they need to use the same airports (turning radii, RWY widths), fit into the same parking envelopes (wingspan) and be serviced by the same ground vehicles (bay heights, wing clearances). Because redesigning an entire industry worth of ancillary equipment and infrastructure has been deemed not worth the minor efficiency gains to be had from flying wings.

• and the extremely conservative attitude of the people making purchasing decisions, that makes it very hard to get even things that look or sound a bit different from the established norm from getting adopted (think the Boeing Sonic Cruiser concept, or the Beechcraft 2000, as prime examples). – jwenting Apr 16 '19 at 5:02

Flying wings simply don't have much internal space for cargo, so they're a non-starter for cargo planes.

You mention the B-2 which will carry 18 tons of bombs. However, bombs are small and heavy: for example, a US Mark 82 bomb is essentially a 130kg (300lb) metal box filled with 90kg (200lb) of explosives. Most airline cargo isn't packed in thick, heavy metal boxes like that, so turning the B-2's bomb bay into a cargo bay wouldn't create a very useful cargo plane.

Which is good, because the designation C-2 is already taken. *rimshot*

I would like to discuss the stability argument in a bit more detail. Since it is correct that static longitudinal stability is the main reason why these aircraft are not often developed.
However the reasoning given in the other posts is incomplete/not completely correct.

First of all, a flying wing indeed has a very small stability margin. This can be solved by either some unconventional wing designs: this has the problem of defeating by large the efficiency gain of using a flying wing configuration.
The other method, employed by the B2 spirit is to use an active controller to control control surfaces. This has the drawback of increasing complexity of the aircraft and passing regulations tests is even harder. some reference.

Static longitudinal stability

I'm going to explain the static longitudinal stability in detail a bit more. First we define stability: to be stable means that whenever a small excitation is applied to the object, the object will "recover" itself.
Longitudinal stability means that an excitation in the longitudinal direction, thus a change in pitch/angle of attack ($$\alpha$$), needs to be countered by "some" moment. Since an aircraft during cruise in equilibrium, an increase in angle of attack, should lead to a negative moment. - A reduction of angle of attack should lead to a positive response moment.

Or in a mathematical way: (definition)

$$\frac{\partial M}{\partial\alpha} < 0$$

A simple wing

Now let us first look at a simple configuration: just a wing. Since lift generated from a wing is due to a distributed force, a wing will always have both a Lifting force, and a lifting moment (except at a single point where the moment is zero, however this point changes with flying conditions). - In aviation we remove the units for simplicity's sake. So we have a force $$C_L$$ and a moment $$C_M$$.

On an airfoil there is also a point where the factor between $$C_L$$ and $$C_M$$ doesn't change with angle of attack. This point is called the aerodynamic center and is a static point given by the airfoil shape: it is hence used to calculate.

So (by definition):

$$\left( \frac{dC_m}{dC_l} = 0 \right)_{a.c.}$$

Now since a wing always generates more lift under a higher angle of attack, and actually we consider the C_L - \alpha curve to be linear. (For stability we consider small changes in angle of attack) the following holds:

$$\frac{d C_L}{d \alpha} = C_{L_\alpha} > 0$$

Together with the earlier equation:

$$\frac{d C_M}{d \alpha} = C_{M_\alpha} > 0$$

conventional aircraft

I first wish to address the stability of conventional aircraft in this point, as there seems to be a lot of contradicting information.

For this consider the following configuration (notice that the points where the lift "attaches" to the wing & tail are defined to be the aerodynamic center for these calculations - we could use any point, but using ac reduces complexity a lot).

From the static equilibrium equations:

$$W = L_W + L_t$$

$$L_W = \frac{1}{2}\rho V^2 S_w \frac{dC_L}{d\alpha}(\alpha - \alpha_0)$$ (above is just the lift equation, which defines $$C_L$$)

The lift due to trim in the tailplane is more complex (due to the non negligible down wash of the main wing on the airflow at the tail ($${\epsilon}$$). ($$C_l$$ = lifting coefficient of tail section)). - Simplifying, we consider the horizontal tailplane to be a symmetric airfoil, so lift at $$\eta=0$$ is zero. (of the tailplane).

$$L_t = \frac{1}{2}\rho V^2 S_t \left( \frac{d C_l}{d \alpha} \left( \alpha - \frac{d \epsilon}{d \alpha} \right) + \frac{d C_l}{d\eta}\eta \right)$$

Similarly the moment equation can be written:

$$M = L_Wx_g - (l_t - x_g) L_t$$

Now from the very first equation again, the partial differential of the moment equation with respect to the angle of attack needs to be negative:

$$\frac{\partial M}{\partial \alpha} = x_g \frac{\partial L_w}{\partial \alpha} - (l_t - x_g) \frac{\partial L_t} {\partial \alpha}$$

Now there is a final definition that needs to be made, a distance $$h$$ from the center of gravity so that for the total wing the moment equation can be written as:

$$M = h(L_w + L_t)$$

Solving all equations (see wikipedia for details) leads to:

$$h = \frac{x_g}{c} - \left( 1 - \frac{\partial\epsilon}{d \alpha} \right) \frac{C_{l_\alpha}}{C_{L_alpha}} \frac{l_t S_t}{c S_w}$$

With $$c$$ being the main aerodynamic chord of the main wing. (Introduced once again to reduce the amount of units we work with). For stability (since $$C_{M_\alpha}$$ needs to be negative) $$h$$ needs to be negative. Let's analyze above result:

$$\frac{l_t S_t}{c S_w} = V_t$$

This part, called the "tail volume", consists of geometric definitions of an aircraft and won't change.

$$1 - \frac{\partial\epsilon}{d \alpha}$$ are the stability derivatives and difficult to calculate, but typically found to be at least $$0.5$$.

So this allows us to define the stability margin as:

$$h = x_g - 0.5cV_t$$

Note that since the second term is always positive, having a negative $$x_g$$, or (see image above) having the center of gravity in front of the aerodynamic center of the main wing. will always give a stable configuration. And remember that aerodynamic center does not change with angle of attack. (Center of gravity can shift during cruise due to fuel consumption, but this is typically mitigated in practice by pumps, and shifting center of gravity forward will always give a more stable aircraft).

neutral point

Now finally we are at the neutral point, which was used in another answer incorrectly consistently. The neutral point is, by definition, the point at which an aircraft is "just" stable: $$h=0$$

$$x_g = 0.5cV_t$$

From this it follows that the "range" between which the center of gravity can change is between nose of the aircraft (negative $$x_g$$) and a point given by mainly the tail volume. The tail volume is most easily influenced by changing either the tail surface or distance between main wing and tail.

Flying wing configuration

Finally back to the original point, the flying wing configuration. A flying wing, by definition, has no tail behind the main wing. Thus the tail volume is zero.

Hence the neutral point of a flying wing is exactly at the aerodynamic center. Which is for a conventional wing design about 1/4th of the chord distance.

thus a flying wing has, without modifications, an unusable small stability margin

Delta wing and canard

I'd also wish to quickly sidestep to the delta wing and canard configuration such as for the concorde or f16. These designs are driven by another parameter (shockwave drag/something else, like more efficient control due to no downwash).

However the stability for such aircraft is a lot different: while the picture above can still be used, we need to consider that $$l_t$$ is, by design, negative. This changes the location of the neutral point to always be in front of the main wing. And many of those designs also have active control surfaces and are inherently unstable.

(The name "canard" even came from this: when the brother wright created the first powered aircraft, in France people didn't believe it. They called it what we would call today "fake news". The term for fake news was "canard" in France, so they called the design a "canard").

• Off to a good start. Now, "tailless" consideration. Make wing chord longer (lower AR). More stable (slower rate of pitch). Now extend a portion of the wing forwards and backwards ( fuselage) even slower pitch. Now flatten the rear portion of the fuselage. (Even more stability). "tailless" flying wings use the trailing edge as a "tail". It is just not as effective as a conventional one for trim when CG is not directly under C all lifts. Hang gliders illustrate this. Weight forward will make an arrow more stable. When a wing is added, CG and Clift imbalance must be trimmed. +1 4 U. – Robert DiGiovanni Apr 17 '19 at 16:24
• Using the trailing edge as a tail doesn't change the fact that for stabilit the tail volume is zero. - This is an aerodynamic property and not a property of the aircraft. It is indeed often done, as I said in the opening paragraph. The effect of this is that the aerodynamic center moves backwards (remember the definition of the aerodynamic center). It's hard to predict the effect without going into CFD though. – paul23 Apr 17 '19 at 18:21
• Trim on the horizontal control surfaces, is however of no influence for the stability, without correct trim settings the aircraft is still (most often) "stable". It's just in a stable slope that either increases or lowers altitude: that is still stable though. (Might not be what you want at that moment, but that's for the pilot to decide not the aircraft design). – paul23 Apr 17 '19 at 18:25
• The "fake news" sense of canard is very recent. Having the control surfaces at the front was called a canard configuration not because it was unbelievable but because it was first used on the Santos-Dumont 14-bis, which was said to look like a duck ("canard", in French) in flight. Also, neither Concorde nor the F-16 has canards. – David Richerby Apr 18 '19 at 19:15
• I'm sorry, I lost the point that you were trying to make. – Koyovis May 26 '19 at 4:52

It's all about CG range and how much abuse the design can take. Take a look at the C-130 Hercules. It has a huge Hstab to cope with a wide range of CG. Really a bi-plane. So is the Chinook helicopter. Holding the table up with 4 legs (6 with a canard).

So, what do we do to get to a viable flying wing? Sweep back offers improvement in pitch stability as (with washout) you lengthen the aircraft. Control surfaces can be placed at the wing tips. Reflexed camber airfoils also help. How to cope the loss of a longer fuselage/Hstab pitch torque arm? Have the cargo bay set on a roller at CG. Pull it forward until it tips. Secure, cargo balanced! Fuel tanks can be arranged to drain evenly. Assuming a subsonic design with near neutral static stability, it may even fly without computers.

But the all important shift in Clift with change in AOA or airspeed must be accounted for. So a small tail, like birds have, may help build a better safety margin for the design, with or without computers. Ditto for lower aspect wings. Interestingly, a bird sweeping its wings back becomes ... a delta. Sweep them back out ... an F-111?

It is possible to reduce tail size in cargo, and passenger planes.

The first certified stable aeroplane flew in 1910 at (and without) the hands of J W Dunne. It was also the first tailless swept plane to fly, sort of a biplane flying wing except that everything was piled in between the wings so not a true flying wing. Contemporaries Handley Page and Igo Etrich were obliged to add tails to their more bird-like attempts. Whether a given tailless type is adequately stable is complex and subtle to analyze and many designers have since got it wrong. In 1913 Dunne lectured the Aeronautical Society in no uncertain terms on why his worked and the others failed, it makes a fascinating read even today.

But all have been agreed that the subsonic tailless type has a narrow CG range. That is not a problem provided you do your load balancing properly, but it does make the chore even more awkward than usual.

The real killer for cargo planes is that the hold of a flying wing only gets deep enough to be practical on a huge design, otherwise the wing would be over-thick and slow. No existing plane has ever been made large enough to make it worthwhile. For it to make sense you need a payload upwards of 500 tons (equivalent to around 5,000+ passengers), six times the Airbus A380 or three An-225 cargo planes or two Stratolaunch Rocs. Oh, and the airports to fly it from.

Simple economics. Why spend billions and years designing a new plane from scratch - especially one that uses technology unproven in civilian applications (flying wing) - when you can spend millions and months buying passenger planes that use proven, tried-and-tested technology, and refit them for cargo needs?

While all the other answers tackle quite a few practical problems that flying wing cargo planes would need to combat, there is also the problem that airplane operators tend to be very conservative when buying expensive aircraft. That's a major reason why commercial airplane design hasn't really changed in the last 50 years. Buying aircraft with a radical new design is risky. Better invest in proven technology that might less efficient rather than to risk loosing your whole investment if the new design turns out to be a failure.