Ok, bear with me: (scroll down for conclusion)

Steady rate of climb (ideal climb situation, real values lie about 25-35% lower) is defined as the power available minus the power required, divided by the aircraft weight.

**(P**_{available}- P_{required}) / W

Available power P_{a} is known, 8 Wasp Major engines of 2640 [kW] each, results in total available power of** 21.12 [MW]**. If you assume a propellor efficiency of around 75%, this results in around **15.8 [MW] of power available.**

The weight W can be guesstimated: assuming an MTOW of 180000 [kg], this is 180000 * 9.81 = **1.77 [MN]**

Required power P_{required} is not known here, but is defined as the drag force D multiplied by the airspeed V. Both of these values are also not directly known, but can be deduced with some assumptions. As wikipedia quotes a cruise speed of around 408 [km/h] or ~ 113 [m/s], a reasonable assumption of climbing velocity seems somewhere around **75 [m/s]**.

The drag D can be calculated by the probably familiar formula **D = C**_{D} * 0.5 * rho * V^{2} * S. The air density **rho **can be taken at sea level to be **1.225 [kg/m**^{3}], the velocity **V **was calculated above and the wing surface area **S **is known to be** 1050 [m**^{2}]. Now only the drag coefficient C_{D} is unknown.

C_{D} is defined as** C**_{D0} + (C_{L}^{2} / (pi * A * e) ). The Aspect ratio **A **is defined as wing area divided by the square of the span, or 1050 / 96.9 = **10.8 [ ]**. Oswald factor **e** is assumed **0.6 [ ]**, since it's not easily calculated and a relatively old design. 0-lift drag coefficient** C**_{D0} is assumed to be **0.025 [ ]**, a rather reasonable value since the aircraft has huge extrenal features like floats...

Lastly, the lift coefficient** C**_{L} needs to be found, which can be done by entering the airfoil data in JAVAFOIL. I plotted the mentioned airfoil, but seeing the shape, I'm having a hard time believing this is what was used. The program gave me a maximum lift coefficient of 1.8, which is rather high, so a more reasonable value of **1.1 [ ]** will be assumed.

This gives a drag force **D** of (0.025 + (1.1^{2} / (pi * 10.8 * 0.6) ) ) * 0.5 * 1.225 * 75^{2} * 1050 = **305 [kN]**.

Therefore, the power required **P**_{required} will be 305 * 10^{3} * 75 = **22.9 [MW]**

This is larger than the power available, so the rate of climb will not be positive. This means the aircraft would not be able to climb, so certainly wouldn't be able to fly. What's bothering me is that this difference seems a little too large for my taste. Since I did make a lot of assumptions here, it could very well be there are some mistakes there...

If you spot any errors or have any theories on why this answer's resulting from this calculation, or have some better values for some of my assumptions, please let me know!