Otdfbt

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)

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.

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*}

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.

I wish to handle 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 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)

$$fracpartial Mpartialalpha < 0$$

A simple wing

Now let us first look at the actual easier situation: 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 easiness. 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):

$$left( fracdC_mdC_l = 0 right)_a.c. $$

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:

$$ fracd C_Ld alpha = C_L_alpha > 0 $$

Give nthis and the equation earlier:

$$ fracd C_Md alpha = C_M_alpha > 0 $$

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:

$$W = L_W + L_t$$

$$L_W = frac12rho V^2 S_w fracdC_Ldalpha(alpha - alpha_0)$$
(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)). - For easiness we consider the horizontal tailplane to be a symetric airfoil, so lift at $eta=0$ is zero. (of the tailplane).

$$L_t = frac12rho V^2 S_t left( fracd C_ld alpha left( alpha - fracd epsilond alpha right) + fracd C_ldetaeta right)$$

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:

$$fracpartial Mpartial alpha = x_g fracpartial L_wpartial alpha - (l_t - x_g) fracpartial L_t partial alpha$$

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:

$$h = fracx_gc - left( 1 - fracpartialepsilond alpha right) fracC_l_alphaC_L_alpha fracl_t S_tc S_w$$

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:

$$fracl_t S_tc S_w = V_t$$

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

$$1 - fracpartialepsilond alpha $$ 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 of the duration of a cruise due to fuel changes, 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 also changed a lot: 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").

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.