Black Horse

Frequently Asked Questions

This FAQ is not intended to cover the basics of the Black Horse proposal; that information is in the other papers on the Black Horse which are on the Black Horse W3 page. Instead, this FAQ is meant to archive questions and answers about details which are not either not present or loosely covered in the other documents, but which have been addressed previously.

[This list was compiled by Daniel Risacher, mostly from postings to the Usenet newsgroup, and is not intended to be comprehensive or authoritative. It is just a collection of things people have said about the Black Horse concept. There is no real order to the questions. Last changes: 2 OCT 95]

Q: The Black Horse is apparently designed to land in an unpowered horizontal mode. Wouldn't a powered horizontal landing be safer? With the fuel tanks almost empty, the fuel consumed while landing should be small.

A: Somebody asked Mitch about this at Space Access 95. His answer was that you don't want even a subset of the engines running at landing time, because the aircraft is too "clean" and it picks up speed rapidly if you light any engines. It might be possible to light an engine or two when you decide you need a go-around.
(Henry Spencer,

Q: Horizontal Lander concepts must drag large mass fractions up and down in the form of wings. A Vertical Lander could very well de-orbit and land with a mass of propellant equal to 10% of the vehicle dry mass; could any HL possibly have wings which are less than 30% of the total dry mass? How many existing aircraft even do that well?

A: This is not correct. The NASA Access to Space baseline SSTO, which you can read about in Aersoapce America, has a landed mass of 251,362 pounds. The wing weighs 11,465 pounds, the tail weighs 1,577 pounds, and the control surface actuation weighs 1,549 pounds. This means that the subsystems that are chargeable to horizontal landing weigh 5.8% of the landed mass. Of that landed mass, 222,582 pounds is the vehicle dry mass, the rest being 25,000 pounds of payload and some residual fluids and such.

For Black Horse, the maximum landed weight is the takeoff weight, which is 48,454 pounds. The wing (1,572 pounds), tails (739 pounds), and control surface actuation (372 pounds) together weigh 5.5% of the landed weight. This aircraft was designed using standard fighter aircraft design methodology, for which see Raymer, Daniel P., "Aircraft Design: A Conceptual Approach" (Chap 15).

For a vertical lander such as the DC-Y, if the landing propellant weighs 10% of the landed mass, that is the equivalent of adding 1200 feet per second to the overall delta-V the vehicle must produce. That means a huge difference at the margin, because the additional delta-V is applied on the steepest part of an exponentially sensitive curve.
(Mitchell Burnside Clapp,

Q: Is the Black Horse a replacement or an alternative for the Shuttle ?

A: Neither; it flies different missions for a different customer in different ways. The current design is meant to be essentially an X-plane with some military utility. Payload to orbit is circa 500kg (plus two crew, who are counted as part of the vehicle, not the payload), but orbital cargo delivery is a secondary role -- the primary military mission of an initial system would be on-demand orbital reconnaissance.
(Henry Spencer,

Black Horse should be able to perform some of the shuttle's missions, and to do so much better than the shuttle. It can also perform missions the shuttle can't, like launching on a few hours' notice to deploy a replacement for a GPS satellite that some future incarnation of Saddam Hussein just destroyed.
(John Schilling,

Q: How much the Black Horse cost per unit? What will the costs of a Black Horse launch/flight be? What is the expected usable life span of each Black Horse space plane? What is the Black Horse's turn around time (i.e. between flights)?

A: I don't have exact numbers on tap, but Mitch has said (if I recall correctly) that he'd be happy, for starters, with the sort of results obtained by the SR-71: a fleet of 5-10 aircraft capable of flying perhaps one mission a day (i.e., individual aircraft turnaround of a few days at most) for a number of years on a total operating budget of under $100M/yr. (This follows usual military practice of treating production and startup costs as sunk costs and not charging them against operations.)
(Henry Spencer,

Q: The papers on the Black Horse all talk about using JP-5 or JP-4 as a propellant, but you certainly can't burn JP-4 in your rocket. This was tried some decades ago, and it was discovered that the variation in composition is just too great. Don't you mean RP-1? RP-1 is essentially the same stuff as JP-4, but with a much tighter spec to avoid burning variation.

A: Actually, the reason one uses RP-1 is to have a coolant-grade hydrocarbon. Ordinary JP-4 or JP-5 will coke in the coolant passages and cause burn-throughs to happen. RP-1 is a low sulfur, low olefin version of JP-4 for that reason. The Russians routinely use ordinary kerosene in their staged combustion engines, at much higher heat fluxes than we tolerate in US designs.

If you use H2O2 as the oxidant, then there is no need to use the hydrocarbon as a coolant at all.
(Mitchell Burnside Clapp,

Q: I thought the point of the Black Horse was to avoid cryogens, since it uses JP-5 and H2O2. In Clapp's papers though, he talks about using LH2 and LOX! Isn't it safer to avoid cryogens completely?

A:Well, the BH design isn't final yet. There are definite trade-offs between LH2/LOX, CH4/LOX, RP/LOX and JP5/H2O2. Currently, the non-cryogenic option seems to be the best for the Black Horse, but there are advantages to the other combinations as well. For instance, there is a better current experience base in LOX handling procedures than H2O2; there are more engine designs available for cryogenic fuels; LOX is cheaper, easier to make, and more stable than H2O2.

Cryogens are slightly dangerous because they are very cold, but the real danger is that they evaporate into clouds of H2 or O2 which can either ignite or speed combustion. On the other hand, H2O2 decomposes spontaneously and exothermically in the precence of impurities. LOX doesn't do this. In the end, just about any rocket fuel is hazardous if handled improperly.

The point of the Black Horse concept is not to avoid cryogens. It is to reduce Delta-v to orbit, simplify operations, and thereby reduce launch costs while increasing response time.
(Daniel Risacher,

Q: How much interest are the people with the money showing ?

A: Nobody is throwing large amounts of money at the program yet. But there is considerable interest. The USAF has long wanted this sort of capability (the buzzword is "trans-atmospheric vehicle") and this is the first design concept that has actually looked realistic and affordable.
(Henry Spencer,

Q: Why do this complicated propellant transfer thing? Why not just drop-launch it from some big jet, like a B-52, or air-launch it off the back of a 747?

A: There are five good reasons to reject the notion of launching Black Horse from another vehicle altogether. First, it's dangerous. You have to make two aircraft fly well when joined, be able to separate safely, and still fly well after the separation. There is no way to build up in flight test to a separation; you either go for the whole thing or you don't try it. It is possible to accept this level of risk, but doing so makes it less attractive and much more expensive. Second, we have limited experience with aircraft-aircraft separations. In aviation history, we've done maybe 400 of them. We did more refueling than that every shift during Desert Storm. There is no complexity associated with in-flight propellant transfer -- it is as easy and as routine as instrument landing. Third, aircraft-aircraft separation reduces the altitude for separation over inflight propellant transfer. All other things being equal, the ensemble is under the thrust of only one aircraft's engines and under the drag of both aircraft, plus the interference drag between them. Fourth, a new facility, perhaps a crane, will be needed to mate and demate the ensemble, which limits basing flexibilty. Finally, the carrier aircraft requirements are much more difficult than for a tanker. It needs to bear not only the weight of the propellant for orbit, but also the dry weight of the orbiter, and its payload, and the mating and separation hardware. This dwarfs the weight of the refueling system. For this reason, you have to develop an entirely new aircraft, or make major, airworthiness-affecting structural modifications to an existing aircraft. Tanking is safer, more familiar, better performing, more flexible, and cheaper than aircraft-aircraft separation.
(Mitchell Burnside Clapp,

Another major reason is that Aerial Propellant Transfer (ATP) can support a much larger payload than an air-launched vehicle. In the case of APT: (a) the aircraft need only carry the oxidizer, and not the weight of the structure, payload, and fuel, and (b) an aircraft can carry a much larger amount of liquid cargo (fuel or oxidizer) than solid cargo, since the former can be carried in the wings. This is why a KC-10 can carry the oxidizer needed for a 9 ton payload, whereas the enourmous AN-225 (by far the worlds largest airplane) is needed to air-launch a 7-8 ton payload. Thus the aircraft support for ATP should be much less expensive than for direct air-launch.
(Larry Gales,

Q: How will the Black Horse protect itself from melting when it re-enters?

A:Reentry heating is a strong function of wing loading. The Space Shuttle has a highly loaded wing, at over 120 lbs/ft2. The Black Horse has a 20 lbs/ft2 wing loading. Some at Boeing believe it could be possible to build an all metal aircraft (Applying Inconel, Rene 41, etc.) since their in-house RASV design used all metal integrated structure/tankage/TPS. (And handled cryogens, too!).
(Mitchell Burnside Clapp,

The feasibility study baselined durable tailored advanced blanket insulation (DuraTABI) material, which weighs 1.1 pounds per square foot, for area ("acreage") coverage and carbon-silica carbide (C/SiC) for the nose, wing, strake, and rudder leading edges. Detailed aerothermodynamic reentry calculations may indicate a less stringent requirement for thermal protection than was assumed in the initial design, possibly even allowing an all-metal skin (Rene 41 or Iconel 617). On the other hand, retaining excess thermal protection - perhaps by applying more advanced thermal protection systems - could give the vehicle a larger reentry envelope and even more operational flexibility.
(The SPACECAST 2020 paper on Space Lift)

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