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Delta III and IV

On 10 May 1995 McDonnell Douglas announced plans to develop the Delta III, a next generation expendable launch vehicle.(1) Building on the success of the Delta II, McDonnell Douglas plans to develop this new intermediate-class rocket with its own funds. The Delta III will be designed and developed at the McDonnell Douglas Aerospace (MDA) facilities in Huntington Beach, CA. The final assembly of the Delta III will take place in Pueblo, CO, with final checkout and launch at Cape Canaveral Air Station, FL.

The payload capacity of the new Delta III will be 8,400 pounds to geosynchronous transfer orbit, more than double the capability of the existing Delta II. The new intermediate-lift rocket will be capable of launching Hughes' largest satellite model, the three-ton, body-stabilized HS 601.

The initial customer for Delta III is Hughes Space and Communications International, a unit of Hughes Electronics Corp., which is owned by General Motors Corp. Hughes and McDonnell Douglas signed a contract for 10 firm launches, plus options for additional launches through 2005. The total value of the contract, depending on options exercised, could be up to $1.5 billion. With an anticipated first launch during the first half of 1998, and the tenth satellite in 2002, options could extend the launches to 2005.

The most significant changes in Delta III's evolution from the existing Delta II are a new single-engine, cryogenically propelled upper stage and a larger fairing to house the payload. The cost to McDAC to do the upgrade is estimated at "more than $200 million."(2)

Major changes to the system include:

New cryogenic upper stage, with a number of details still uncertain, although the company claims to have a detailed design in hand. The engine has yet to be identified, with the company suggesting candidate engines from Aerojet, Rocketdyne and Pratt & Whitney. The current betting suggests the Centaur's RL-10 or derivative (which would seem to provide about the right performance). In general, this new cryogenic upper stage would seem to be the single largest area of risk in the program -- particularly in light of the fact that every other cryogenic upper stage has exhibited sognificant development problems. McDAC Program Director Rick Arvesen has acknowledged that this will be the greatest hurdle for Delta III.(3) Thus it may be anticipated that McDAC will take a rather conservative design approach.

Expanded 13.1 foot diameter fuel tank for the first stage (the LOX tank retains the existing 8 foot diameter). By growing the diameter of at least part of the core stage, the vehicle can accomodate more propellant, while keeping the existing length (which reduces controllability problems).

New 13.1 diameter payload shroud.

Upgraded solid motors (46" diameter versus 40", 10 % longer), with three of the nine motors featuring thrust vector control (currently none of the SRMs have TVC). The need for TVC on the solids appears driven by concerns that the existing core engine of the Delta II would not provide sufficient control authority for the new configuration. These would be provided either by Hercules or Thiokol, selected through a competitive procurement.

New avionics and software, although the details have not been enumerated in published reports. As with past Delta upgrades, the avionics system will probably be modified to match the new configuration, with control system software and hardware modified to support the new solids and vehicle configuration.

Although the new cryogenic upper stage resembles a single-engine Centaur, with the purchase of General Dynamics by Martin Maritta, it is difficult to imagine LockMart providing this produce to a competitor. But, in many respects the new vehicle resembles an Americanized H-2. This may clarify how McDAC will remain viable in the the EELV competition -- increasing performance and reducing development costs by proposing Americanized versions of the H-2. This would make considerable sense from McDAC's side, since they otherwise would have had considerable difficulty matching LockMart's EELV. At a minimum, the comparison of the H-2 and the Delta III presents a strong family resemblance.

One key design element supporting this conjecture is the 13.1 foot diameter shroud and tankage for the new upper stage -- the H-2 has a constant 13.1 foot diameter tankage and shroud. The H-2 second stage and the Delta III diameter's seem too close to be coincidence. The diameter for the Delta III (in feet) is exactly 4.00 meters -- which might be overlooked, given the tendency to assume that when an American company quotes something in feet, that they are therefore actually working in feet. In this case, it appears McDAC is the first American company to be building a metric rocket.

McDAC has been partnered with the Japanese on N-1/H-1, so it only makes sense that now the design/hardware flow would be reversed. This also would make sense for the Japanese, since they have invested non-trivial sums into H-2 with very little prospect of commercial return. The H-1 was a Delta derivative, arising from the licensed production of the Delta 1 as the N-1, but the H-2 was a completely in-Japan development. This was driven partially by Japanese government policy, but also to allow them to potentially compete in the international launch market (the licensing agreements for the Delta-related technologies did not allow them to produce similar systems with American-based technologies for commercial markets).

It seems unlikely that McDAC would simply buy outright H-2 second stages, and attach another engine.

From an economics standpoint, recent months have witnessed one of the sharpest appreciations of the yen relative to the dollar in recent history. Thus it is difficult to imagine that McDAC would choose this time to agree to buy increasingly expensive Japanese hardware (particularly given the risk that it would become even more expensive in the future).

A more likely possibility is that McDAC and MHI (Mitsubishi Heavy Industries) reverse roles from the past. Previously McDAC has facilitated the growth of the modern Japanese aerospace industry through licensed production by MHI of aircraft like the F-4 and F-15, and launchers such as the Delta. The pattern has usually been that the Japanese first bought one or two whole vehicles for local assembly for initial familiarization, gradually moving to domestic production on their own assembly lines.

However, there are several counter arguements suggesting that the Japanese connection is more apparent than real:

The core stages are quite different since the H-2 uses a cryogenic LOX/LH2 core to a Mitsubishi LE-7 engine, and the Delta-III uses a LOX/RP stage to a single Rocketdyne RS-27A engine. Different fuels, different engines, different design.

The strap ons are also dratically different, and their usage will probably be different too, if MDC sticks to their staggered ignition and discard approach on the Delta II. H-2 has two large segmented strap-ons which run the full length of the LH2 core tank, and which each have gimbal systems.

Assuming the Delta III is just a scale up of the Delta II, the Delta III solids will use 9 monolithic solid motors, with only 3 having TVC capability. From the view-graphs, they look like they tie their loads in about at the top of the LOX tank (aft cg versus the forward cg on the H-2). If they use the Hercules design as the scale up, they will have a wound composite case, instead of the H-2 bolted steel case.

Initial reporting on the Delta III implied that McDAC had not selected the second stage engine, though it was implied that the H-2s LE-5 was not among the candidates. There may be a Japanese political reason for that: the Japanese basic law that established NASDA prohibits development of anything that is later used for non-peaceful purposes. This has been interpreted to allow civilian earth resources satellites, but to prohibit military reconnaissance or similar functions. Thus a Delta III using H-2 hardware would almost certainly not be allowed to be used in the EELV program. Perhaps the H-2 s tankage might be used, since it may be debatable whether the H-2 tankage is considered as "NASDA developed", because it was not totally new technology. But the H-2's LE-5 and -7 engines were new NASDA-sponsored developments. In practice, this might depends on which Japanese political party (Socialists or Liberal Democrats) is in power when the question is decided.

A number of lines of reasoning also counter the apparent connection between the cryogenic upper stages:

From a viewgraph perspective one cryo stage tends to looks much like another. But the second stage on the H-2 uses a common bulkhead design for the combined LOX/LH2 tankage -- whereas the Delta III uses a separate LOX and LH2 tank system according to the published sketches (this is a more conservative design approach, keeping with what I think would be the MDC approach to minimize cost and technical risk in the new cryo stage design.

The 4.0 meter diameter shroud is a fairly standard size (available not only on H-2, but also Ariane 4). The fact that a US launch vehicle company decides to make a payload shroud that is four meters is not a conclusive element of proof. McDAC is fully capable of designing large payload fairings (they make the extremely large Titan 4 payload fairing, for example). The fact that the Delta 3 and H-2 fairings look similar can also be easily explained by the common design heritage which MHI inherited from MDC, not the reverse.

The 13.1-foot diameter tankage for the new first stage may simply be derived from the diameter of the payload shroud. By avoiding the "hammer head configuration" you avoid some potentially problemsome aerodynamic and accoustic effects. And, from a cost perspective McDAC might be able to reuse some of the tooling needed for the shroud to manufacture the core stage (or at least pieces of the core stage).

The H-II second stage may be too large (wet mass = 19.7tons) for a Delta. given that the current Delta 2nd and 3rd stages only weight 7 and 2 metric tons, respectively, raising questions as to whether the 1st stage core could support that much extra weight. Also, the H-II second stage was designed to carry fairly heavy payloads in the 10t+ range. It is unclear whether Delta III will launch anything that large. Consequently, the upper stage would be over-designed for the job - the mass ratio is (3t/19.7t)=0.153 vs. 0.1-0.11 for Centaur-A/T & Ariane.

It seems unlikely that McDAC is going to buy outright H-2 second stages, and

attach another engine. On the other hand, it may be the case that McDAC and MHI (Mitsubishi Heavy Industries) reverse roles from the past. The absence of MCDAC saying that they are in partnership with H-2 is also puzzling, which at a minimum suggests that the negotiations are still in progress. It would appear that MHI has been prevented from either selling or licensing its LE-5 engines to MDC and others because of the restrictive interpretation regarding peaceful purposes (i.e. MDC could not use an MHI upper stage on a vehicle used for US military launches). While this would not prohibit MDC from selling in the commercial market, it would be unlikely that they would readily cede the potential economies of scale that could be achieved if they could address the US military market as well.

This second stage may be MHIs attempt to get something out of the H-2 program by going in with McDAC. Things were looking depressing enough for the Japanese company trying to commercialize the H-2 to begin with; and the recent surge in the yen had to really depress them. Now, a truly commercial competitor, right in their own market range, comes to the fore, with 10 solid orders to kick them off. Thus it may be that the ambiguity on this point reflects the incomplete state of negotiating the terms of the agreement. The companies may still be trying to figure out respective roles & missions, particularly with respect to the EELV US content requirements. If MHI gets a commission each time Delta-3 flies, they stand to gain revenue from commercial operations they would otherwise never see.

Delta 4 -- EELV

McDonnell Douglas is evaluating four Russian engines for potential inclusion in its EELV proposal:(4)

NK-33 was designed by the Samara State Scientific and Production Enterprise (Kuznetsov) during the 1960s as part of the N-1 lunar landing rocket, and would be licensed for production in the United States by Aerojet. The N-1 first stage used 30 NK-33 engines, each with 1,510kN of thrust. The second stage used eight of the similar NK43 engines, which had larger nozzle. Aerojet has plans to begin manufacturing the NK-33 in the US within three years of a contract award, and during May 1995 brought two of the engines to the US for testing (at its own expense).(5)

RD-170 designed by NPO Energomash of Moscow (for potential production under license in the US by Pratt & Whitney) is a four-chamber engine that powers the Zenit launchers, and served as the strap-on booster for the Energia/Buran. The RD-170 runs at some 24,000 kPa (3,560 psi). The LOX/RP-1 staged combustion engine uses two 14,500-rpm turbopumps to drive two sets of paired thrust chambers to full thrust at sea level of up to 7,903 kN (11.77 million Ibst.).

RD-180 designed by NPO Energomash of Moscow is a two thrust chamber version of the four-chamber RD-170. About 80% of the RD-180 is common to the RD-170 and other Russian designs. Overall the RD-180 is 157 inches long, with a diameter of 118 inches, including frame, sensors, hydraulic system, and gimbal actuators. The nominal thrust sea level is 8,277 kN (900,000 lbst), with a specific impulse at sea level of 309 seconds, and vacuum ISP of 337 seconds. The RD-180 can be throttled over a 40-100% thrust range

RD-190 designed by NPO Energomash of Moscow is a single bell derivative of the LOX/RP-1 RD-170. The boost pumps, turbopumps, preburner, preburner fuel regulator, and fuel throttle are scaled from the RD-170, while other components (main combustion chamber, gimbal actuators, hot gas ducts, health monitoring system and injectors) are the same as those of the RD-170. With a sea-level thrust of 430,000 pounds and a vacuum thrust of 430,000 pounds, the specific impulse of this engine varies from 311 seconds at sea level to 337 seconds in vacuum. Capable of throttling from 40% to 100% of rated power, the engine operates at a chamber pressure of approximately 3,560 psia. The RD-190 is 144 inches long and 70 inches in diameter, and weighs 5,548 pounds.

Although the company has not publicly revealed how these engines might be integrated into its EELV proposal, it would seem likely that they would be added to a 4 meter diameter core stage, with tankage derived from the H-2 booster. This might produce a "Delta 4A" which could also add two or three additional upgraded (46" diameter, 10 % longer) solid motors for additional payload capacity. Satisfying the EELV heavy lift requirements might neccesitate a "Delta 4B" with two 120" diameter solid rocket motors (perhaps derived from the Hercules Solid Rocket Motor Upgrade SRMU program), or a "Delta 4C" with four such motors might be required (these vehicles represent such an excursion from existing Delta configurations that precise calculation of their performance would be required to determine the neccessary configuration to meet EELV requirements).


1. C-reuters@clarinet.com (Reuters), buenneke@rand.org (Dick Buenneke), z9405597@cc.mita.keio.ac.jp (David Paul Huntsman), kingdon@cygnus.com (Jim Kingdon), larrison@ix.netcom.com (Larrison), mlindroo@news.abo.fi (Marcus Lindroos INF), alastair@firewall.ihs.com (Alastair Mayer), mark.oderman@channel1.com (Mark Oderman), "Delta-3," sci.space.policy, May 1995.

2. Wall Street Journal, 8 May 1995.

3. "Matching Delta II's Success May Be Tough for MDC's New Launcher," Space Business News, 17 May 1995, page 3

4. Ferster, Warren, "US EELV Effort Eyes Soviet-Origin Engines," Space News, 8 May 1995, page 10.

5. Bill Sweetman and J.R. Wilson, "Getting There," International Defense Review, January 1995, pages 29-33.

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