Von Braun selected nitric acid/hydrazine propellants, perhaps as a result of the same Peenemuende research that influenced the French team. He very conservatively used a combustion chamber pressure of 15 atmospheres (as in the V-2 engine). Ideal engine performance for each stage was based on combustion chemistry and gas dynamics calculations. The nozzle expansion ratio differed for each stage, so 'based on experience' Von Braun allowed for expected efficiency losses of 5%, 7%, and 10% on the first, second, and third stages. Total nozzle exit area for each stage was calculated, but no particular detailed engine design was presented. However the height of the engine nozzles was indicated for vehicle sizing.
Specific impulse and thrust for the engines was not expressed in common modern terms, which causes some confusion in the literature. The figures given are the specific impulse at the condition of the ambient atmospheric pressure equalling the nozzle exit pressure. The figures adjusted to vacuum and sea-level specific impulse terms indicate the upper stage versions of the engines would have a specific impulse of 297 seconds, a very respectable value not achieved in flight for a storable propellant engine until 1962.
The nitric acid/hydrazine propellant combination used was found not to be practical without additives. These were developed during the course of the 1950's. But by 1962 nitric acids were replaced as a storable oxidiser by nitrogen tetroxide. Furthermore, in comparison with Von Braun's 15 atmosphere assumption, by 1962 chamber pressures of 40 atmospheres had routinely been achieved, allowing higher expansion ratios (and higher performances) for first-stage engines.
Von Braun used a constant thrust/weight value of 69:1 in calculating the engine mass for all stages. In fact this would be nearly met on the next US large engine design after the V-2, the Redstone. Von Braun's engine used a separate propellant (hydrogen peroxide) to drive the engine pumps. This was the practice used on the V-2, and in Russia was continued even in the R-7 ICBM (which still flies today as the Soyuz launch vehicle). But thereafter the US and Russia moved quickly to turbopumps driven by the same propellants used in the same engine.
The 1948 calculations assumed the exhaust nozzle was in the shape of a continuous annular ring around the central parachute container in the first and second stages. The nozzle completely occupied the base of the third stage. The concept almost presages the aerospike engine designs of the 1960's. It may have been a continuation of the V-2 philosophy of having many combustion chambers feed a single nozzle. It may have been that this was just a means for roughly sizing the base of the launch vehicle.
In the 1952 Collier's series the number of engines per stage were indicated, allowing their mass, diameter and thrust to be allocated. However it must be noted that these diameters cannot be reconciled with the geometry of either the 1948 or 1952 stage designs! Nevertheless it is interesting to compare the figures calculated by Von Braun for his engines with the actual figures achieved for the Titan 2 ICBM engines, the only US storable propellant engines of this size ever developed, first flown in 1962:
|Parameter||Von Braun, 1948||Titan 2, 1962|
|Isp (sea level)-sec||217||259|
When considering Von Braun's estimates for the structural weight of his rocket, he arrived at values near those achieved in the late 1960's for the Saturn and N1 vehicles in the US and Russia. His values look larger at first, but they include the mass of recovery equipment (parachutes and soft-landing retrorockets) that amounted to 18% and 11% of the empty weights of the first and second stages respectively. Deducting these amounts, and the hydrogen peroxide consumed during the ascent, and it is apparent that Von Braun calculated values for the structure of his launch vehicles essentially equivalent to those for the only vehicles of those size built in the US and Russian in the late 1960's:
|Structure||Von Braun 1962||Saturn V 1968||N1 1969|
|Stage||First Stage||S-IC||N1 Block A|
|Stage||Second Stage||S-IB||N1 Block B|
Third-Stage Re-entry Vehicle
The third stage itself made a long re-entry glide for 22,000 km from atmosphere entry interface at 80 km -- 55% of the earth's circumference! During this time the glider experienced a peak deceleration of 0.45 G and a peak airframe temperature of 1005 deg K. The airframe of the glider was seen as reaching a thermal equilibrium during the long glide, where it could re-radiate heat as quickly as it was absorbed. Based on this Von Braun concluded the glider could be built of existing steels (although he admitted that 'latest information' indicated the peak heating could be 300 deg K higher).
Research during the 1950's showed that the higher value was nearer the truth and that the temperatures at the leading edges of re-entry vehicles far exceeded the theoretical average values calculated in 1948. The Dynasoar design of the early 1960's, the closest analogue to Von Braun's design to reach an advanced stage of development, would have experienced peak nose temperatures of 2140 deg K, with the structure stabilising at 1255 deg K during an equivalent long-glide re-entry. It was necessary to build the Dynasoar's external skin from coated molybdenum, and the internal structure from Rene-41 nickel alloy. Von Braun's plan to use conventional steel would not have worked. Furthermore, to cool the pilot and payload, it was necessary to encase the cockpit and payload bay in a heavy 'water wall' insulation/cooling system. The lightweight refrigeration system used in Von Braun's calculations would not have been adequate.
The conventional solution for re-entry design became the use of ablative or exotic heat-radiative skin materials, and to get the re-entry accomplished as quickly as possible to avoid heat-soaking the structure. The space shuttle, for example, experiences actual maximum external temperatures of 1800-1250 deg K, lands just 5450 km down-range from an altitude of 80 km, and pulls a maximum of 2 G's during re-entry. The internal structure is nearly completely protected from the heat of re-entry (as long as there is no breach, as on STS-107), and is primarily 2024-T81 aluminium alloy.
However a minority of the technical community continues to advocate Von Braun's long-glide, high-lift, re-radiative approach. These included Nonweiler waverider advocates, who proposed to actively cool the sharp, hot leading edge of their designs but claimed the rest of the spacecraft could be built from conventional metal structure. The following compares several designs to show the effect of wing loading on peak and minimum underbelly heating for gliding re-entry vehicles.
|Peak heating deg K||Min underbelly deg K|
|Von Braun 1948||73||>1400||1100-1400|
|Black Horse 1996||102||1640|
The Recovery Systems
Von Braun's glider was designed to deliver nominally 25 tonnes of cargo, plus 14.5 tonnes of 'excess propellant'. This would be used to assemble the enormous planned Mars expedition - 70 crew aboard ten spacecraft with a mass of 3720 tonnes each!
Since assembly of an expedition of this size would take 950 launches, Von Braun was very concerned to make sure all of the launch vehicle was recoverable and reusable. The first stage would return to earth under a 64.5 m diameter parachute weighing 85 tonnes! Just prior to splashdown 40 tonnes worth of solid rockets would ignite and allow the stage to splashdown at zero velocity 304 km downrange. It would be towed back to the Johnston Island launch site by a tug, refurbished, and reused. The total mass of the recovery provisions amounted to 18% of the stage empty mass. The second stage required a 20.4 diameter chute with a mass of 3.8 tonnes and 4 tonnes of braking rockets. It would be recovered 1459 km down-range and the recovery provisions were 11% of the stage empty mass.
If we delete the recovery provisions and the manned glider provisions from the design, we can calculate that the three stages Von Braun's design could deliver 60 tonnes, or one percent of its lift-off mass, into a 200 km polar orbit. By comparison, the three-stage Titan 3A (with transtage) could deliver 1.33% of its lift-off mass into the same orbit. This again emphasizes the modernity of the design.
LEO Payload: 25,000 kg (55,000 lb) to a 1,730 km orbit at 23.50 degrees.
Stage Data - Von Braun 1948
Status: Study 1952.
Gross mass: 6,400,000 kg (14,100,000 lb).
Payload: 25,000 kg (55,000 lb).
Height: 97.00 m (318.00 ft).
Diameter: 20.00 m (65.00 ft).
Thrust: 119,300.00 kN (26,819,700 lbf).
Apogee: 1,730 km (1,070 mi).