Prepared Statement by Michael Griffin 8 May 2003 (part 2)


Continued from part 1

- Given that the OSP program has not yet progressed beyond establishing the Level I requirements, do you think NASA's plan for spending approximately $750 million on technology demonstrations between FY03 and FY06 is justified? What technologies are the most critical to demonstrate before proceeding to full-scale development?

Numerous advances in thermal protection materials technology have been made since the Shuttle was designed and built, and some relatively inexpensive demonstrations may be useful in this area. Automated rendezvous and docking, a procedure so basically straightforward that the Russians first demonstrated it more than three decades ago, remains to be demonstrated in the U.S. program. Crew escape system technology has been essentially absent from U.S. vehicles since Apollo, and may need some investment. Isolated technology demonstrations may be required to address issues relevant to a particular vehicle design, once such a design is selected. However, these are details. I am unaware of any crucial, but as yet unproven, technology needed for Earth-to-LEO transportation. I believe money spent on technology demonstrations would, in general, be better spent on vehicle development. Such an approach would also offer the benefit of significantly shortening the planned OSP development schedule.

- What design alternatives should NASA examine as it performs its concept studies for the OSP? What changes to the OSP program would you recommend to reduce the cost or accelerate the schedule? How does the decision to proceed with a design that is totally reusable, partially reusable, or expendable drive design complexity, development schedule, cost, and safety? Can the OSP schedule be accelerated significantly without introducing unwarranted risks? If so, what recommendations do you have?

We should be careful to avoid overburdening OSP with ISS crew return vehicle (CRV) requirements. My view, harkening back to my involvement in the 1993 Space Station redesign effort, and before, has always been that the CRV is properly viewed as a "lifeboat", to be used in an emergency, and likely not otherwise. As an order of magnitude estimate, we might expect to use it once per decade. If it is used regularly or routinely, we are doing something seriously wrong with regard to the operation of ISS, something which needs to be remedied. But stretching the notion of what constitutes a CRV is not the answer. Therefore, again in my view, crew transport requirements should determine the OSP design, with CRV requirements at the margin.

As an aside, I have personally never been able to understand why a refurbished Apollo spacecraft cannot be outfitted as a perfectly acceptable CRV. The need for developing a new vehicle to meet the crew escape requirement has never been obvious to me.

Much in the news recently, and for good reason, is the question as to whether the "Orbital Space Plane" should be a "plane" at all. In the wake of the Columbia disaster, some have called for a return to a "capsule" design, more properly termed a "semiballistic entry vehicle". Certainly there is strong merit in such a recommendation. A semiballistic vehicle offers a number of advantages for Earth-to-LEO transport. It is likely to be more volumetrically efficient and to have less mass than a winged vehicle for the same overall mission requirements, and is much better adapted to any requirements to go beyond low Earth orbit. Either design can be equally reusable, with the possible exception of the heat shield for the semiballistic vehicle, which will almost surely encounter a higher heat load than for a gliding entry vehicle. However, and in strong contrast to a winged vehicle, the semiballistic can be designed such that the heat shield is both very simple, completely separable, and easily detachable from the core vehicle, resulting in a system with only one non-reusable component that is not particularly weight critical and can be, almost literally, dirt cheap.

It is often stated that the landing accuracy of a semiballistic vehicle will be inferior to that of a winged design. This is nonsensical. If a parachute or parasail is used, today's steerable designs, with pinpoint GPS guidance, allow either design to achieve highly accurate landing point control. Furthermore, historical data indicates that even without benefit of steerable parachutes and GPS, entirely acceptable landing accuracy can be obtained. The table below cites the mission-by-mission Apollo landing accuracy (from "Apollo Program Summary Report", NASA TM-X-68725, National Aeronautics and Space Administration, Johnson Space Center, Houston, TX, April 1975). It is seen that the worst-case landing dispersion would have been trivially contained within the boundaries of Edwards AFB, or White Sands Missile Range, or even within acceptable landing areas at Cape Canaveral or Wallops Flight Facility. Most of the Apollo landing dispersions would have fitted easily within the boundaries of Dulles Airport. It is not necessary to do better than that.

Apollo Landing Accuracy

MissionDistance
from Target(mi.)1
Apollo 71.9
Apollo 81.4
Apollo 92.7
Apollo 101.3
Apollo 111.7
Apollo 122.0
Apollo 131.0
Apollo 140.6
Apollo 151.0
Apollo 163.0
Apollo 171.0

1 Best estimate based upon recovery ship positioning accuracy, command module computer data, and trajectory reconstruction.

Note the phrase above, "if a parachute is used". It is not obvious that a parachute is necessary (other than possibly as a backup system, wherein the goal becomes crew, rather than vehicle, survival). The terminal velocity of a semiballistic vehicle will be on the order of 300 miles per hour, probably less. Braking rockets ignited at high altitude, initially at idle thrust, and then smoothly throttled to touchdown can serve quite well, as the DC-X and DC-X-A programs have shown. Besides demonstrating the ultimate in pinpoint landings in the nominal case, these efforts also showed how a backup parachute landing system can be efficiently incorporated into the design, and used effectively in an emergency. Detailed studies have continued to reveal no substantive mass difference between a semiballistic design with terminal rocket braking, and a more traditional winged design.

Of course, there is also the possibility of using conventional parachute descent, with surface contact cushioned by short-duration, high-thrust rockets as in the Soyuz design. Thus, there is no need to assume the inconvenience of an Apollo-style water landing if a semiballistic design is chosen, except possibly in a dire emergency when, in contrast to a winged vehicle, the ability of a semiballistic to survive a ditching then becomes an attractive option.

However, because we should carefully consider the merits of a semiballistic crew vehicle design does not mean that we should ignore the merits of a winged design. Various lifting body research programs, as well 198 successful X-15 flights and 116 successful Shuttle landings (including approach and landing tests with the Enterprise vehicle) have demonstrated the efficacy with which unpowered descent and landing can be performed. Highly efficient blended delta-wing, lifting body shapes, such as the NASA Langley HL-20 and its derivatives, have been thoroughly characterized. So there is a wide range of attractive options available.

When considering winged vehicle designs, however, I think we have ignored one of the best options, the straight-winged design, for somewhat specious reasons. All else being equal, it is well understood that a straight-wing design will have less mass, lower heat loads, a higher subsonic lift/drag ratio, a lower landing speed, a shallower glide path on approach, and better subsonic handling characteristics than a comparable delta-wing design. The delta-wing design offers as its principal advantage a somewhat greater entry crossrange capability than for a comparable straight-wing design. This allows greater maneuverability from orbit to reach a given landing site, as opposed to waiting on-orbit for perhaps half a day for another opportunity to reach the site. The delta-wing design also allows the so-called "abort once around", meaning that the Shuttle can land at its launch site after only one orbit, in the event of a severe anomaly. This greater atmospheric maneuverability was the reason for its selection for the Space Shuttle design, and was a source of considerable controversy at the time. But in over a hundred Shuttle flights, operational practice has shown that this enhanced crossrange capability is at most a minor convenience, rather than a significant enabling feature. Any consideration of a new, winged, spaceplane should take these facts into account in determining a design configuration.

When contemplating designs for a new winged space plane, it may not be beyond the bounds of reason to examine the swing-wing concept, so successful on the F-14 fighter aircraft. Providing robust, mass-efficient thermal protection of the wing leading edges is among the most difficult, and unforgiving, tasks in a spaceplane design. With a swing-wing concept, it might be possible to avoid this task altogether. For such a vehicle, the atmospheric entry phase would be performed as a semiballistic design, while terminal area energy management, approach, and landing would be performed as a conventional winged vehicle. As always, there are tradeoff analyses to be conducted, but the concept may be worth pursuing.

The issue of OSP reusability is complex, which of course is why it attracts so much debate. The primary reason to prefer a reusable vehicle is that, in all reason, it should be cheaper to operate. Secondary reasons may include the fact that ground and flight crews gain experience with the nuances of a particular machine, a valuable benefit when compared to the obvious risks of undertaking a maiden voyage for every flight of an expendable vehicle. However, for the moment let us restrict the discussion to economic issues.

The economic benefits of reusability are strongly conditioned by the cost of incorporating the necessary features into the design and fabrication of the vehicle, and by its assumed flight rate and operational lifetime. As a simple example, if it will cost five times more to build a reusable vehicle than to build a comparable expendable design, the reusable vehicle must fly five times to break even with the expendable, assuming their processing costs are similar. Moreover, most of the cost for the reusable vehicle is incurred "up front", while a greater proportion of the expendable vehicle cost is incurred only when the next unit is actually procured. Time-value-of-money considerations can thus strongly benefit the expendable vehicle when flight rates are low, and when decisions are made on a lowest-life-cycle-cost basis.

The issue of designing to minimize life-cycle cost is worth some discussion. It should be noted that, over more than two decades of Shuttle operation, the program has encountered much criticism because year-to-year operational costs have been quite high when considered on a per-flight basis. This has been directly traced, in part, to early-1970s budget constraints on initial design and development, when numerous choices were made which had the effect of minimizing (or appearing to minimize) development cost, while increasing operational costs. Again because of time-value-of-money considerations, the strategy of designing the vehicle to minimize development cost is closely akin to that of a design based on minimizing life-cycle cost, especially when the vehicle will be in service for a long time. While neither principle is inherently wrong, each should be applied in moderation. Life cycle costs are heavily biased by early-year, or "up front", costs. It is always easy to defer operational funding problems to the "out years". Yet, when the "out years" arrive, as they always do, we seem consistently to regret the pattern of earlier choices, which were of course intended to "save" money. Is it possible, this time, that we could at least make a new mistake?

As outlined earlier, it will be tempting on economic grounds to consider an expendable design for OSP, for the reasons just mentioned. I believe this is a mistake; if done, it will represent a failure of government to lead where industry, by itself, cannot go. An argument to go backward, toward deliberate use of expendable vehicles for manned spaceflight, is an argument which inevitably favors the doing of less manned spaceflight, precisely because out-year operational costs will always been seen as unacceptably high when the out-years arrive. This should not be our goal.

With respect to cost, I would like to offer a cursory figure of merit, a target cost-per-pound of delivered hardware. It is well established within the aerospace community that such figures of merit offer a valid first-order estimate of likely program cost; indeed, such parameters form the basis of all accepted cost models. Therefore, I would advocate that the OSP design, development, test, and evaluation (DDT&E) costs should be upper bounded at $100,000 per pound for the dry mass of the vehicle. The nation's experience base with reusable manned space vehicles is limited, but both X-15 and the Space Shuttle orbiter would seem to fit this definition. In recent-year dollars, both were completed at a DDT&E cost of approximately $90,000 per pound of delivered hardware. If the OSP is allowed to cost more, we are conveying the message that nothing at all has been learned in 40+ years of manned spaceflight.

Regarding the program schedule, it seems inconceivable to me that a nation which required only eight years to reach the moon, from virtually a standing start, can require a similar or greater length of time to design and deploy a simple crew transport vehicle. If the OSP program requires more than five years - at the outside - from authorization to proceed until first flight, it is being done wrong. My primary recommendation, the only one I think can affect the outcome in a significant manner, is this: Define carefully the goals the OSP is to meet. Pick a strong, effective, proven, and trusted program manager, and accord to him or her the total authority and responsibility for success. Set aside the necessary funds, with adequate margin. And then see to it that everyone else stays out of the way.

- What challenges may NASA face in using an Expendable Launch Vehicle (ELV) as the boost vehicle for the OSP? Does the use of an ELV for human spaceflight pose an unacceptable risk?

In the 1950s and 1960s, the term "man rating" was coined to describe the process of converting the military Redstone, Atlas, and Titan II vehicles to the requirements of manned spaceflight. This involved a number of factors such as pogo suppression, structural stiffening, and other details not particularly germane to today's expendable vehicles. The concept of "man rating" in this sense is, I believe, no longer very relevant.

If a winged design is chosen for OSP, there will be an issue of coupling between the OSP vehicle aerodynamics and the launch vehicle structural dynamics. Briefly, the OSP must be oriented and flown very close to its zero-lift aerodynamic angle of attack. Any significant amount of lift on the OSP wings will create lateral loads at the OSP/launch vehicle interface that are quite likely unacceptable, at least without additional structural reinforcement at that interface. However, it must be said that launch vehicle loads are likely not the limiting factor; the wings of a spaceplane cannot themselves accept high lateral loads without being ripped off. The problem is a familiar one; the Shuttle must be flown with a nearly zero angle of attack for similar reasons.

Therefore, irrespective of the launch system used for a winged OSP, the vehicle must be flown at essentially a zero-lift angle of attack, and any variations due to vehicle aeroelasticity must be carefully controlled. While the problem is certainly not trivial, it is not likely to be any more difficult for the new evolved expendable launch vehicle (EELV) than it will be for a winged OSP attached to a future RLV.

The base reliability of unmanned expendable vehicles seems to arouse concerns where that of the manned Shuttle system inexplicably does not. Many, if not most, unmanned payloads are of very high value, both for the importance of their mission, as well as in simple economic terms. The relevant question may be posed quite simplistically: What, precisely, are the precautions that we would take to safeguard a human crew that we would deliberately omit when launching, say, a billion-dollar Mars Exploration Rover (MER) mission? The answer is, of course, "none". While we appropriately value human life very highly, the investment we make in most unmanned missions is quite sufficient to capture our full attention.

Logically, therefore, launch system reliability is treated by all parties as a priority of the highest order, irrespective of the nature of the payload, manned or unmanned. While there is no EELV flight experience as yet, these modern versions of the Atlas and Delta should be as inherently reliable as their predecessors. Their specified design reliability is 98%, a value typical of that demonstrated by the best expendable vehicles. If this is achieved, and I believe that it will be, and given a separate escape system with an assumed reliability of even 90%, the fatal accident rate would be 1 in 500 launches, substantially better than for the Shuttle. Thus, I believe that launching OSP on an expendable vehicle would pose no greater risk - and quite likely somewhat less risk - for human spaceflight than is already accepted for the Shuttle.

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