Douglas ROMBUS

In 1964, Phil Bono of Douglas Aircraft Co. proposed a low-cost heavy lift VTVL SSTO RLV plus lunar base as a logical follow-on to the Apollo project. Bono's "Reusable Orbital Module-Booster & Utility Shuttle" (ROMBUS) concept was based on his patented plug nozzle rocket engine design, which doubles as a heat shield during atmospheric reentry. The vehicle's base-first reentry mode assures a stable condition during recovery since the mass of the engine is very high, i.e. far aft center of gravity. The plug nozzle would be cooled by circulating liquid hydrogen through the same regenerative system used for cooling the engines and base of the vehicle while the engines are operating during ascent. Mixture ratio of liquid oxygen to hydrogen was raised to 7:1 -- about as close as one can get to the stoichiometric value of 8:1 without running into combustion chamber cooling problem. During ascent, the plug nozzle provides automatic altitude compensation and therefore good performance at both sea level atmospheric pressure and in space. For final orbital insertion, 16 of the 36 engines would burn for 3 seconds to provide the required velocity. ROMBUS would typically spend 24 hours in orbit before the ground track passes close enough to the launch site for de-orbit. Parachutes and (beginning at 0.73km altitude-) retrorockets would be used to safely land the vehicle. The final touchdown burn would be provided by four engines running at 25% thrust for approximately twelve seconds.

To reduce the size and weight of the vehicle, the hydrogen fuel was to be stored in eight external jettisonable tanks. The tanks were jettisoned and then recovered by parachute as they were depleted during ascent to orbit. The earlier fully reusable "ROOST" concept was thus rejected since it did not leave much margin for growth in vehicle structure mass.

The total life cycle cost would have been $10 billion [1964 $'s] over 10 years including $4.088 billion for the development program. Bono mentioned the following SSTO RLV-specific advantages : increased reliability since each vehicle has a history, reduced development cost & complexity vs. multi-stage vehicles, economies of scale possible since plug nozzle engine units & tanks could be mass produced. The estimated direct launch cost was $22.4 million (=$28/lb. to a 568km orbit at 1964 economic conditions) and the planned vehicle turnaround time about 76 days. Bono also mentioned a direct operations cost goal of $12/lb (5-6 reuses) - $5/lb (>20 reuses) for a vehicle payload capability of 450t by the year 1975. In comparison, the Saturn V was then expected to cost $150-250/lb. The vehicle would have used the same Kennedy Space Center facilities as the Saturn boosters, although a new launch pad would have been required.

Liftoff Thrust: 80.1 MN. Total Mass: 6,350.3t. Total Length: 29m.

Payload capability: 362,274kg to a 568km Earth orbit.

Development cost: $22 billion at 1999 economic conditions. Launch cost goal of $60-$330kg [1999 $s].

Stage Number 1: ROMBUS. 36 x plug-nozzle engines (3000psia pressure, 1:7 mixture ratio). Gross Mass: 6,350.3t incl. ~5,000t LOX. Empty Mass (core vehicle only): 306.175t. External fuel tank mass: (18.143t + 89.358t of LH2 fuel) * 8. Liftoff Thrust: 80.1 MN. Isp: 455s [vac.]. Length: 29m. Width: 24.4m. Propellants: LOX/LH2.

Sources:

"ROMBUS concept" -- Bono, Astronautics & Aeronautics 1964/January/p.28

"All-Purpose Moon Vehicle" -- Space World 1964/December/p.16

"Reusable Booster for Logistics and Planetary Exploration" -- Bono,Woodworth & Ursini, IAF XV:th INTERNATIONAL ASTRONAUTICAL CONGRESS PROCEEDINGS 1964/p.547

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ROMBUS could fly all the way to the lunar surface and back if low Earth orbit (LEO) rendezvous & propellant transfer were used. The fully loaded LH2 tanks would be transferred from from ROMBUS tankers to the lunar landing ROMBUS, hence the need for pumping low-density hydrogen is circumvented. The tankers would also carry oxygen in their payload compartment which is pumped to the lunar vehicle tank. Phil Bono felt ROMBUS might be able to compete on a cost basis with traditional specially developed expendable spacecraft such as Apollo, although the total launch weight would be higher since no weight-saving boosters were used.

Second illustration depicts an expendable ROMBUS landing on the Moon. Only two LEO tanker flights would be required for soft-landing a 4,536kg payload on the lunar surface. The empty spacecraft has a mass of 306.233t and the structure + tankage might be transformed into a lunar habitat.

Nine LEO tanker missions would be required to land a 226.8t payload on the lunar surface and return the vehicle to Earth with 22.68t of cargo. With seven tanker flights, the lunar payload capability would be 45.359t plus 4.536t of return payload. After ROMBUS has landed on the Moon, its LH2 tanks (now empty) could be transformed into habitation modules for a moonbase. Each 7-tanker mission would leave two fuel tanks behind while the 9-tanker flights would leave four tanks. ROMBUS would have to carry more LH2 coolant (+5,670kg) to survive the high-velocity atmospheric reentry. But this difference would still be far smaller than the 313t of propellant required to propulsively cancel the 3,048m/s difference between reentry from lunar transfer orbit and reentry from low Earth orbit. The Douglas engineers did not expect the thermodynamics to cause significant problems, although they did note that the point of maximum heating on the ROMBUS plug-nozzle underside is different during lunar return: some additional LH2 might be diverted by suitable valving to cope with this problem. The underside of the vehicle would also have to be beefed up to cope with the additional dynamic pressure during reentry from lunar transfer orbit.

Bono's Project Selena called for the establishment of a 1000-man lunar colony in four successive phases by 1984. The main purpose of the Selena base was to support three unmanned Mars cargo delivery missions by 1986 and about half of the payloads (16,029t by mass) would be dedicated to the "Deimos" Mars follow-on. The total mass of the lunar cargo was a staggering 32,950 metric tons, requiring 1341 ROMBUS launches and 1011 cargo/propellant transfer operations in low Earth orbit over 8.5 years. A fleet of 10-15 ROMBUS vehicles would have to perform 330 lunar landing missions to deliver the crews, cargo and propellant. Bono lists the following design assumptions for Project Selena:

# 40-flight average ROMBUS lifetime.
# Personnel delivered to Moon at 90.7kg/man
# Life support requirement of 4,536g/man/day
# Maximum lunar tour of duty of 4 months
# Lunar refueling base will store liquid oxygen and hydrogen propellants for Mars mission
# Cryogenic propellant stored in cylindrical tanks at permanently shadowed location on Moon
# Reusable Moon-Mars orbit-Earth vehicle lands 21,772kg on Mars (18,143kg useful payload) and consumes 5,343t of propellants per trip. The fourth, manned, Project Deimos mission then departs from Earth orbit in 1986.

Bibliography:
"All-Purpose Moon Vehicle" -- Space World 1964/December/p.16

"Reusable Booster for Logistics and Planetary Exploration" -- Bono,Woodworth & Ursini, IAF XV:th INTERNATIONAL ASTRONAUTICAL CONGRESS PROCEEDINGS 1964/p.547

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Phil Bono's second major objective was the manned exploration of Mars. Although an Apollo-type expendable spacecraft was considered, Bono felt a modified ROMBUS would be cheaper and safer although the less optimal single-stage design inevitably would translate to a 3.5 x higher launch mass in Earth orbit. The Mars-bound vehicle would weigh 3965.758t in its 323-kilometer parking orbit. This illustration depicts the end of the trans-Mars injection (TMI) burn in May 1986. The four TMI propellant tanks (now empty) are jettisoned.

After an interplanetary cruise period of 200 days, the ROMBUS spacecraft fires its engines to enter a 555-km parking orbit around Mars. The mothercraft now has a mass of 984.75t after jettisoning two more empty fuel tanks. A conical 25-tonne Mars Excursion Module (MEM) then departs with a landing crew of three astronauts.

The MEM's relatively cramped crew quarters and life support systems would only permit a relatively limited 20-day stay on the martian surface. However, future missions would spend up to a year on Mars thanks to the unmanned cargo delivered by previous unmanned flights.

ITEM MASS (kg)
USEFUL PAYLOAD RETURNED TO MARS ORBIT 2,495
M.E.M. ASCENT STAGE (DRY) 1,361
ASCENT PROPELLANT (ISP=426s, v=4572m/s) 7,620
M.E.M. LANDING STAGE (DRY) 6,758
SURFACE PAYLOAD 3,629
LANDING PROPELLANT (ISP=426s, v=549m/s) 3,084
TOTAL 24,947
Mars Excursion Module Weight Summary

The crew would return to the waiting mothercraft in an 11,475-kg Mars ascent vehicle that separates from the empty MEM landing module. Its RL-10 rocket engines consume 7620kg of oxygen & hydrogen propellant during the 4572m/s ascent burn. Finally, 280 days after entering Mars orbit, the ROMBUS spacecraft fires its engines to return to Earth. The remaining two fuel tanks are then jettisoned after the final 3749m/s trans-Earth injection maneuver. The empty 340,194 kg vehicle lands on Earth 330 days later. The total round-trip travel time is 830 days.

The "Deimos" mission would naturally have required some modifications to the basic ROMBUS design, including the addition of two SNAP-8 nuclear reactors for auxiliary power. The 6-crew would have lived inside a toroidal Command Center module which also housed 6300kg of food, oxygen & water (1.27kg/day/person).

DATE MANEUVER MASS (kg) DELTA-V (m/s)
ROMBUS GROSS WT. IN 323km EARTH ORBIT 3,965,737 3,658
05/09/86 VEHICLE AT MARS ESCAPE 1,747,682
11/25/86 VEHICLE ENTERS 555km MARS ORBIT 984,744 2,347
09/21/87 PREPARE FOR TRANS-EARTH INJECTION 963,879
09/21/87 VEHICLE AT MARS ESCAPE 415,035 3,749
08/16/88 VEHICLE AT EARTH ENTRY (12192m/s AEROBRAKING) 382,830 119
08/16/88 VEHICLE LANDS ON EARTH SURFACE 340,193 171


Bibliography:
"Reusable Booster for Logistics and Planetary Exploration" -- Bono,Woodworth & Ursini, IAF XV:th INTERNATIONAL ASTRONAUTICAL CONGRESS PROCEEDINGS 1964/p.547

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The aerospace transportation industry grew by leaps and bounds during the 1960s as bigger, faster and ever more capable aircraft and rockets became available. There appeared to be no limits to progress. In April 1963, NASA's Manned Spaceflight Center released a launch forecast for "Nova"-class heavy rockets in the 500-tonne payload class. The expected need for such vehicles in 1975-90 included a 50-crew Moonbase (1975-), manned Mars flights (1981-), unmanned planetary craft (1979-), large space stations (1980-),and a "global transportation system" (1980) with 242 flights/year by 1990, and a "Nova military strike force" in 1976 (15 flights/year by 1981). Douglas Missile & Space Systems tried to capitalize on this by proposing a series of suborbital rockets capable of transporting 110-260 passengers at 25,000km/h. The "Hyperion" vehicle was truly remarkable since it would have been launched horizontally and landed vertically (HTVL) -- an extremely rare combination. The payload capability was 110 passengers or 18t of cargo. The takeoff mode was similar to contemporary HTHL TSTOs, I.e. a subsonic sled riding on a cushion of air. Hyperion would be travelling at 1100km/h as it leaves the sled at the end of the 3km launch rail. Unlike other Douglas SSTO concepts, Hyperion was fully reusable so it would have been ideally suited for flights from inland sites since no fuel tanks would be dropped during flight. The booster sled would literally have provided a "flying start" which greatly reduced the SSTO dry mass. Unfortunately, the Hyperion launch system also required a 1.7km high mountain so Douglas mostly regarded the concept as an experimental vehicle.

Douglas expected that the vehicle would cost $1.5 billion to develop (=$8 billion at 1999 economic conditions), and the cost per seat would have been $3000 -- or $15000-16000 in 1999.

Bibliography:
Frontiers of Space -- Phil Bono and Kenneth Gatland, 1969.

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The "Pegasus" was a Saturn V-class intercontinental rocket capable of transporting 170-260 passengers and 13-33.5t of cargo at 25,000km/h, or 90t to a 560km low Earth orbit. It would have been available in the 1980s and reduced the travel time from New York to Bombay from 22 hours to only 40 minutes. In Phil Bono's words: "the SST would be just a small step in the direction of reducing transit time, whereas rocket-propelled vehicles would provide the ultimate in sub-orbital speed for transportation on the Earth's surface. The space age has equipped us with the technology for transporting emergency equipment to any disaster area on Earth in the time it would take for a bus ride across town, despite the conditions of aircraft runways upon arrival. In addition, diplomats, heads of state, chiefs of staff, as well as business executives, could be transported in person to consult, confer, inspect, decide and to lead in a manner currently impossible from distant shores, even with the aid from television."

Pegasus would have carried most of its hydrogen fuel in expendable drop tanks; hardly advisable for a low-cost suborbital launcher but necessary in order to improve the maneuvrability of the rocket during reentry and landing since the vehicle then could be made much lighter and smaller. The projected gross liftoff mass was 1520t and the landing weight only 148t. This illustration depicts the emergy landing procedure, where all eight tanks are jettisoned before the vehicle makes an emergency landing on inflatable pontoons. Pegasus would have carried sixteen 1170KN thrust aerospike rocket modules so the engine-out safety would have been fairly good since only four engines were required for landing. The expected landing accuracy was only 1.6 * 3.2 kilometers so the spaceport would have to be located in a 5km wide uninhabited area for safe launch and landing operations.

Pegasus atmospheric reentry. The combined heatshield + plug nozzle rocket engine would be cooled by circulating liquid hydrogen fuel at the base of the vehicle. The intercontinental passenger version would incorporate fins between the external liquid hydrogen tanks to restrict reentry decelerations to a safe maximum of 2.5-3 g's. Pegasus would have been launched from spaceports which, by the 1980s, were expected to be established near most key cities throughout the world. Off-shore launch platforms were another possibility. Douglas estimated that the Pegasus would cost $3 billion [1964 $s -- $16.1B at 1999 rates] to develop, and the operational vehicles would cost $34M (=$183M in 1999 )per copy if two dozen were built. The first operational Pegasus would have cost $63 million (=$338M in 1999) to produce. According to Phil Bono, "we cannot afford to dismiss the rocket transport as a far-fetched impractical pipe dream. We must design today as if the next 10 years had already passed."

Bibliography:
"Pegasus" -- Space World 1964/December/p.20

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Douglas also proposed a military VTVL SSTO for transporting troops and cargo -- the "Ithacus." The Ithacus plan was apparently inspired by general Wallace M. Greene, who proposed a rapid-strike batallion of 1200 soldiers. These "rocket commandos" would have been deployed by an intercontinental rocket so the need for overseas US Army bases might be greatly reduced. The gigantic 6400-tonne "Ithacus Sr." was supposed to serve this strategic need. The basic configuration was very similar to ROMBUS, i.e. eight hydrogen drop tanks + rocket-powered vertical landing using a plug-nozzle aerospike engine. The landed weight would have been about 500t. One of the biggest problem was returning the 64m tall vehicle back to the launch site. Self-ferry was out of the question since the rocket would have produced a liftoff thrust of 80200KN -- way too high without a custom-built launch pad with sound suppression, water cooling etc.. Shorter "hops" of a few hundred kilometers would however have been possible so the plan was to fly the partially fueled vehicle to a coastal harbor. The illustration depicts Ithacus Sr. being transported by a Saturn-type "crawler" onto a barge for return to the United States.

The smaller "Ithacus Jr." version depicted below would have had an intercontinental cargo capability of 33.5t or 260 soldiers. Douglas proposed to launch two Ithacus Jr. vehicles from an Enterprise-class nuclear aircraft carrier, which also would have produced liquid oxygen and hydrogen propellant from seawater. Power for the electrolysis process would have been taken from the carrier's nuclear reactor: 112MW would have been required to produce 1150t of oxygen and 164t of hydrogen from 1470t of water. The rocketships would be stored inside hangars. One Ithacus Jr. would serve as a troop carrier while the other would deploy unmanned cargo to the same military site.

The Ithacus Jr. vehicles would land 600 meters apart to deploy a fully armed group of 260 soldiers.

Frontiers of Space -- Phil Bono and Ken Gatland, 1969

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In late 1966, the vertical launch & landing SSTO proponents at Douglas Aircraft Co. carried out a study to determine whether ballistic VTVLs might be cost-competitive vs. winged VTHL TSTO vehicles in the small payload class. Previous NASA & USAF studies had generally assumed ballistic single-stage vehicles might make sense for unmanned heavy-lift payloads but winged TSTOs were invariably chosen for small manned near-term missions. Consequently, Douglas had to define a small VTVL SSTO manned "space taxi" to demonstrate the key elements of the concept (aerospike engine, lightweight structures, ballistic reentry, vertical landing, actively cooled heatshield etc.) The resulting vehicle became known as "Saturn Application Single Stage to Orbit". Notable design features included an aft-mounted liquid oxygen tank to reduce the difference between vehicle center of gravity & center of aerodynamic pressure, and a hydrogen cooling system for the main engine to provide thermal protection during reentry. Thermal analysis indicated that although the engine itself would be adequately protected by this system, the areas located above the exhaust nozzles would not. Consequently, the designers had to resort to an ablative, expendable material (200 kilograms of Armstrong Insulcork 2760) bonded to the aluminium structure although it would increase the maintenance cost. The oxygen/hydrogen mixture ratio was 6:1 rather than 7:1 since the designers felt a high oxygen ratio would degrade the exhaust velocity & payload capability. 50% hydrogen slush was used to reduce the volume of the fuel tank. The 36-segment plug nozzle propulsion system would have operated at a pressure of 1500psia. It would be used for ascent, orbit insertion, de-orbit and (beginning at an altitude of 760 meters-) the final landing burn. The vehicle would carry enough propellant for hovering for 10 seconds before landing at an unprepared site, if necessary. The estimated landing accuracy of 1853 * 3700 m was not regarded as a major concern since the Gemini 6-12 flights achieved an average touchdown dispersion of only 6.85km although the capsule had essentially no maneuvering capability below 30.5km altitude. The reentry crossrange capability was about +-370km, permitting a safe landing at El Paso, TX or Wendover Range, UT after 2-3 orbits from Cape Canaveral. Wendover was the preferred emergency landing site since SASSTO easily could have been returned from nearby Hill AFB to Cape Canaveral in a "Pregnant Guppy" S-IV-B transport aircraft.

SASSTO had a payload capability of 3,629kg to a 185km orbit and the standard payload would be a 2-man Gemini spacecraft protected by a jettisonable fairing to reduce drag losses during ascent. This would provide a safe emergency escape system for the test pilots, and the Gemini ejection seats, heatshield, paratchutes etc. (1542kg in all) could later be removed as the flight test program increases confidence in SASSTO reliability. Douglas envisioned this vehicle as a "space fighter" capable of satellite inspection missions, or space station resupply flights lasting a maximum of 48 hours. It could also deliver 2,812kg of liquid hydrogen to a spacecraft in Earth orbit.

ince SASSTO was loosely based on the Saturn S-IV-B rocket stage, Douglas also proposed an expendable version for use as a more capable upper stage with the Saturn IB and Saturn V launch vehicles. The expendable SASSTO stage would have had a burnout mass of 7,400kg and carried 85,729kg of oxygen+hydrogen propellant. The stage was thus of a much more lightweight construction than the standard S-IV-B (12,949kg + 104,326kg LOX,LH2) and the new aerospike engine would have been more efficient as well (464s specific impulse vs. 426s for the J-2 engine). Consequently, the Saturn V's payload capability would have been boosted by 8-11t as well. The Saturn IB's basic 15876-kilogram payload capability to a 185km orbit would have increased to 23814-25855kg depending on whether SASSTO would be flown in expendable or reusable mode. The latter version was known as SARRA (Saturn Application Retrieval and Rescue Apparatus) and was intended for returning stranded Apollo crews from the lunar surface.

Finally, the Douglas design team also compared the cost of SASSTO with two different all-rocket VTHL TSTOs: a winged 1st stage plus lifting-body 2nd stage (center) and winged first & second stages (right). All three vehicles were designed for a 2,812-kilogram payload although the lifting-body TSTO only was able to carry 2,086kg due to center of gravity problems. No attempt was made to estimate the marginal launch cost since there were too many unknown factors. VTVL SSTO would however be expected to yield a significant operational advantage since only a single vehicle must be maintained and the VTVL SSTO does not require a landing runway. SASSTO was expected to cost $1.1. billion to develop (=$5.88B at 1999 rates). The winged VTHL TSTO would cost 2.2 times as much to develop as SASSTO while the smaller lifting-body TSTO variant would be 50% more expensive. The winged and lifting-body 1st unit production costs would be 4 and 2.7 times higher than the SASSTO 1st unit cost, respectively. The general conclusion was that the complex winged or lifting-body TSTO shapes result in added liftoff and manufactured weights of a more expensive construction than ballistic wingless SSTOs. For example, the lifting-body TSTO dry mass (12,274kg + 2,086kg payload) is 2.4 times higher, and the winged TSTO weighs 3.6 times as much (18,176kg+2,812kg P/L) as SASSTO at touchdown. The gross liftoff weights bear the relationships of 1.0 (SASSTO; 97,887kg GLOW), 1.25 (lifting body orbiter TSTO; 122,245kg GLOW) and 1.91 (wing-body orbiter TSTO; 187,020kg GLOW). In that case, is the combination of lower reentry g-loads, better maneuvrability (landing go-around with jet engines) and improved crossrange really worth the cost of carrying wings...? Although TSTO thus appears to be uncompetitive vs. ballistic single-stage RLVs for small payloads, the authors admit that requirements for higher payloads (22.68-45.6t) may yield rapid increases in propellant mass fraction for winged two-stage vehicles, making TSTO more performance/cost-effective.

Liftoff Thrust: 1,232.655KN. Total Mass: 97,976kg. Total Length: 18.8m.

Payload capability: 3,674kg to a 185km low Earth orbit.

Stage Number 1: SASSTO. 36 x plug-nozzle engines (1500psia pressure, 1:6 mixture ratio). Gross Mass: 97,976kg. Empty Mass (core vehicle only): 6,668kg. Thrust: 1,232.65-1,557.5KN. Isp=367-464s. Length:18.8m. Width: 6.6m. Propellants: LOX/slush LH2.

Bibliography:

"Enigma of Booster Recovery - Ballistic or Winged? -- Bono,Senator & Garcia, SAE Conference Proceedings 1967/0382/ p.57

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blackstar said:

That's amusing. The artist obviously did not realize that for Voyager, the spacecraft in the illustration was not what it would look like landed on Mars.

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ACEMANN said:

I placed an order for the Pegasus docs a while back - I received them in PDF form on a CD but since I was the first person to request the documents electronically (!!!!!) they had to scan them and that took a while. Has a ton of information on the vehicle. If I was a good 3D artist I'd whip up a model and fly it in orbiter.

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Michel Van said:

SENSITIVE BUT UNCLASSIFIED (SBU)

Wat its that CENSORED for a CENSORED ?

Is that CENSORED NASA way of a CENSORED CENSORED totally CENSORED Copyright ?!

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Orionblamblam said:

Umm.... not really sure what your specific question is here.

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martinbayer said:

Michel,

it's not quite true that NASA ignored Bono's concepts for decades. In fact, then NASA administrator James Webb was granted two US patents for a Recoverable Single Stage Spacecraft Booster "with respect to an invention of Philip Bono", US Patent D201773 from July 27, 1965 for the configuration and US Patent 3295790 from January 3, 1967 for the detailed design, which clearly represent the ROMBUS:
Martin

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Michel Van said:

NASA has the Patents Rights on ROMBUS !

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Orionblamblam said:

Woo! Was just informed that the documents have been reviewed and are no longer ITAR controlled.

V2N6 will have a whole lot of goodies.

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Triton said:

Orionblamblam said:

Woo! Was just informed that the documents have been reviewed and are no longer ITAR controlled.

V2N6 will have a whole lot of goodies.

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Great news! It seems absurd that declassified technical reports are available under the Freedom of Information Act (FOIA) and then they run afoul of International Traffic in Arms Regulations (ITAR).

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Orionblamblam said:

Woo! Was just informed that the documents have been reviewed and are no longer ITAR controlled.

V2N6 will have a whole lot of goodies.

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ACEMANN said:

So what did you end up getting? Pretty much everything MSFC had on it??

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Triton said:

Orionblamblam said:

Woo! Was just informed that the documents have been reviewed and are no longer ITAR controlled.

V2N6 will have a whole lot of goodies.

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Great news! It seems absurd that declassified technical reports are available under the Freedom of Information Act (FOIA) and then they run afoul of International Traffic in Arms Regulations (ITAR).

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