KJ_Lesnick
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Wow, it looks a LOT like the L-1011's cockpit! Thanks Skybolt!
KJ Lesnick
KJ Lesnick
US Supersonic Transport (SST) Program 1960-1971
- Thread starter Orionblamblam
- Start date
Cabin layout for 16 first class plus 108 economy class seats. 1 - passenger doors: 2 - doors for ground staff: 3 - buffet; 4 - clothes hanging space; 5 - washroom; 6 - supply compartment; 7 - cockpit; 8 - seats for stewardesses.
Interavia 10/1961 – Page 1,378
NAA Proposal for a Supersonic Transport
Technical Data
Powerplant
6 ducted fan engines
Thrust with secondary air-stream reheat
31,000 lb per engine
Thrust without secondary air-stream reheat
25,300 lb per engine
Main Dimensions
Span
104 ft
Length
198 ft
Top of fin from ground .
31 ft
Wing
Area
7,000 sq.ft.
Max. wing loading
61 lb/sq.ft.
Mean aerodynamic chord
75 ft
Aspect ratio
1.81
Leading edge sweep
65.67°
Dihedral
0°
Flaps
490 sq.ft.
Control Surfaces
Canard Area
410 sq.ft.
Sweep (25% chord). . .
38°
Dihedral
0°
Flaps
54 sq.ft.
Vertical fins (two) Area .
2 x 255 sq.ft.
Sweep (25% chord)
45°
Fuselage
Max. width (overall)
12 ft. 4 in.
Max height (overall)
9 ft. 2 in.
Cabin length (incl. subsidiary compartments)
89 ft 8 in
Cabin Width
11 ft
Cabin Height
7 ft
Gangway width (1st Class)
23 in
Gangway width (Economy Class)
15 in.
Cargo compartment Capacity
1,000 cu.ft.
Weights
Total structure.....
122,800 lb
Total power plant . . .
50,800 lb
Total equipment ....
30,400 lb
Manufacturing weight
204,000 lb (1)
Max. takeoff grossweight
439,000 lb
Max. landing weight . .
300,000 lb
Max. zero fuel weight. .
260,000 lb
Max. usable fuel ....
217,000 lb (2)
Performance
Cruise speed equivalent to
Mach 3
at average altitude of
68,000 ft.
Critical field length
6,200 ft. (3)
7,900 ft (4)
Landing runway required
7,950 ft. (5)
8,600 ft. (6)
1.) For the 108 economy class and 16 first class seat version. For variations in cabin class 27 lb per economy class and 32 lb per first class seat must be allowed.
2.) 32,400 US Gal. at 6.7 lb per gal.
3.) According to SR 422B; max. takeoff gross weight; sea level; 100°F.
4.) As 3, but 3,000 ft above sea level.
5) According to SR 422B; 40 percent reverse thrust on three engines during ground roll; maximum landing weight; sea level.
6) As 5, but 3,000 ft above sea level.
As an example of the various projects for a Mach 3 transport prepared by American constructors, we have selected here that of North American, concerning which some details were recently announced.
As North American stresses, this is only a paper study, which is subject to continuous alteration. The first efforts of the NAA engineers arc being directed towards reducing takeoff weight by light construction methods and aerodynamic improvements, in order to cut down fuel consumption and reduce operating costs. Later modifications should also enable economic employment of the aircraft on shorter routes at low supersonic, or even subsonic, speeds. Despite their provisional nature, the data available concerning NAA's preliminary design are of interest and are suitable as a basis for operational investigations and commercial calculations.
As is to be expected, this design is inspired (it is not alone in this) by the B-70 Valkyrie Mach 3 bomber. but is fully tuned to civil requirements. The calculated operating costs are comparable to those of today's long-haul jets for 3,500 mile block distances. The run-ways normally available at international airports are adequate for takeoff and landing. And, thanks to the "quiet" but high thrust turbofans, climb performance is good, and there will be little reason for those living in the vicinity of airports to complain of excessive noise. If sound pressure at ground level is not to exceed 1.46 lb/sq. ft, the aircraft must not pierce the sound barrier below 34,000 ft.
Features of Design
North American selects the following features of its design as important:
1. Trimmable canard control surfaces.
2. Spoilers and air brakes.
3 .Downward folding wing tips for economic cruise flight.
4 .Braised stainless steel honeycomb sandwich and riveted titanium sheet construction.
5. Six duct-burning turbofan engines with thrust reversers. 6. Multi-shock induction system —one to each three engines.
7.. Main undercarriage legs each with 8-wheel bogie, tire. pressure 160 p.s.i.
8. 4,000 p.s.i. hydraulic system with brazed fittings.
9., Fully powered automatic hydraulic control system.
10.. Closed-cycle air-conditioning system for the cabin, with ram air pressurization for emergency use.
11. 115/200 volt, 400 cycle, 3 phase AC electrical system.
12 System checkout during flight.
13 Crew vision angles in accordance with current US regulations (CAR 4b).
14. Self-contained hoists for facilitating loading and unloading of 224 in x 110 in .x 76 in freight compartment (usable cargo space 1,000 cu. ft).
Flight Profile
The assumptions used in the calculation of times, speeds and fuel consumption are as follows: climb at subsonic speed to about 43,000 ft; transonic accelera-tion in climb to supersonic speed; attainment of Mach 3 at minimum cruise altitude; continuous climb at Mach 3; average cruise altitude about 70,000 ft; deceleration to subsonic speed at an altitude of at least 43.000 ft; approach flight and landing; plus six minutes manoeuvre allowance. For a block distance of 3,400 n.m. to be accom-plished with maximum payload, resultant block time, (which includes 15 minutes taxiing time), is 2 hrs 45 mins, block speed 1,240 knots, and block fuel consumption (during flight plus 5.000 lb in taxiing) 176,000 lb.
If the aircraft is forced to deviate to an alternative airport, the amount of fuel used at most favourable cruise altitude, at an all-up weight of 300,000 lb and under ISA conditions, is about 1,050 lb per min. If holding at the airport of destination becomes necessary, the fuel requirement at most favourable Mach no. holding flight, with an all-up weight of 250,000 lb at an altitude of 25,000 ft is about 400 lb per minute, and at sea level some 490 lb.
Typical Operational Weight Calculation
This example is based on a block distance of 3,400 n.m., e.g. New York/Zurich, with max. payload. Cabin layout: 16 first class and 108 tourist class passengers.
lb
Manufacturing Weight Empty
204,000
Crew (8 Persons plus Baggage)
1,580
Unusuable Fuel and Oil
1,400
Cooling Water
2,000
Engine Oil
285
Liquid Nitrogen Fuel System
400
Flight supplies, emergency equipment etc
1,945
Operating weight empty (rounded values)
212,000
Payload (Passengers, baggage, freight)
31,000
Operating Zero fuel Weight
243,000
Reserve fuel (for 300 n.m. diversion at cruise altitude plus 30 minutes hold at sea level
25,000
Operating Landing Weight
268,000
Flight Fuel
171.000
Maximum all-up weight
439,000
Interavia 10/1961 – Page 1,378
NAA Proposal for a Supersonic Transport
Technical Data
Powerplant
6 ducted fan engines
Thrust with secondary air-stream reheat
31,000 lb per engine
Thrust without secondary air-stream reheat
25,300 lb per engine
Main Dimensions
Span
104 ft
Length
198 ft
Top of fin from ground .
31 ft
Wing
Area
7,000 sq.ft.
Max. wing loading
61 lb/sq.ft.
Mean aerodynamic chord
75 ft
Aspect ratio
1.81
Leading edge sweep
65.67°
Dihedral
0°
Flaps
490 sq.ft.
Control Surfaces
Canard Area
410 sq.ft.
Sweep (25% chord). . .
38°
Dihedral
0°
Flaps
54 sq.ft.
Vertical fins (two) Area .
2 x 255 sq.ft.
Sweep (25% chord)
45°
Fuselage
Max. width (overall)
12 ft. 4 in.
Max height (overall)
9 ft. 2 in.
Cabin length (incl. subsidiary compartments)
89 ft 8 in
Cabin Width
11 ft
Cabin Height
7 ft
Gangway width (1st Class)
23 in
Gangway width (Economy Class)
15 in.
Cargo compartment Capacity
1,000 cu.ft.
Weights
Total structure.....
122,800 lb
Total power plant . . .
50,800 lb
Total equipment ....
30,400 lb
Manufacturing weight
204,000 lb (1)
Max. takeoff grossweight
439,000 lb
Max. landing weight . .
300,000 lb
Max. zero fuel weight. .
260,000 lb
Max. usable fuel ....
217,000 lb (2)
Performance
Cruise speed equivalent to
Mach 3
at average altitude of
68,000 ft.
Critical field length
6,200 ft. (3)
7,900 ft (4)
Landing runway required
7,950 ft. (5)
8,600 ft. (6)
1.) For the 108 economy class and 16 first class seat version. For variations in cabin class 27 lb per economy class and 32 lb per first class seat must be allowed.
2.) 32,400 US Gal. at 6.7 lb per gal.
3.) According to SR 422B; max. takeoff gross weight; sea level; 100°F.
4.) As 3, but 3,000 ft above sea level.
5) According to SR 422B; 40 percent reverse thrust on three engines during ground roll; maximum landing weight; sea level.
6) As 5, but 3,000 ft above sea level.
As an example of the various projects for a Mach 3 transport prepared by American constructors, we have selected here that of North American, concerning which some details were recently announced.
As North American stresses, this is only a paper study, which is subject to continuous alteration. The first efforts of the NAA engineers arc being directed towards reducing takeoff weight by light construction methods and aerodynamic improvements, in order to cut down fuel consumption and reduce operating costs. Later modifications should also enable economic employment of the aircraft on shorter routes at low supersonic, or even subsonic, speeds. Despite their provisional nature, the data available concerning NAA's preliminary design are of interest and are suitable as a basis for operational investigations and commercial calculations.
As is to be expected, this design is inspired (it is not alone in this) by the B-70 Valkyrie Mach 3 bomber. but is fully tuned to civil requirements. The calculated operating costs are comparable to those of today's long-haul jets for 3,500 mile block distances. The run-ways normally available at international airports are adequate for takeoff and landing. And, thanks to the "quiet" but high thrust turbofans, climb performance is good, and there will be little reason for those living in the vicinity of airports to complain of excessive noise. If sound pressure at ground level is not to exceed 1.46 lb/sq. ft, the aircraft must not pierce the sound barrier below 34,000 ft.
Features of Design
North American selects the following features of its design as important:
1. Trimmable canard control surfaces.
2. Spoilers and air brakes.
3 .Downward folding wing tips for economic cruise flight.
4 .Braised stainless steel honeycomb sandwich and riveted titanium sheet construction.
5. Six duct-burning turbofan engines with thrust reversers. 6. Multi-shock induction system —one to each three engines.
7.. Main undercarriage legs each with 8-wheel bogie, tire. pressure 160 p.s.i.
8. 4,000 p.s.i. hydraulic system with brazed fittings.
9., Fully powered automatic hydraulic control system.
10.. Closed-cycle air-conditioning system for the cabin, with ram air pressurization for emergency use.
11. 115/200 volt, 400 cycle, 3 phase AC electrical system.
12 System checkout during flight.
13 Crew vision angles in accordance with current US regulations (CAR 4b).
14. Self-contained hoists for facilitating loading and unloading of 224 in x 110 in .x 76 in freight compartment (usable cargo space 1,000 cu. ft).
Flight Profile
The assumptions used in the calculation of times, speeds and fuel consumption are as follows: climb at subsonic speed to about 43,000 ft; transonic accelera-tion in climb to supersonic speed; attainment of Mach 3 at minimum cruise altitude; continuous climb at Mach 3; average cruise altitude about 70,000 ft; deceleration to subsonic speed at an altitude of at least 43.000 ft; approach flight and landing; plus six minutes manoeuvre allowance. For a block distance of 3,400 n.m. to be accom-plished with maximum payload, resultant block time, (which includes 15 minutes taxiing time), is 2 hrs 45 mins, block speed 1,240 knots, and block fuel consumption (during flight plus 5.000 lb in taxiing) 176,000 lb.
If the aircraft is forced to deviate to an alternative airport, the amount of fuel used at most favourable cruise altitude, at an all-up weight of 300,000 lb and under ISA conditions, is about 1,050 lb per min. If holding at the airport of destination becomes necessary, the fuel requirement at most favourable Mach no. holding flight, with an all-up weight of 250,000 lb at an altitude of 25,000 ft is about 400 lb per minute, and at sea level some 490 lb.
Typical Operational Weight Calculation
This example is based on a block distance of 3,400 n.m., e.g. New York/Zurich, with max. payload. Cabin layout: 16 first class and 108 tourist class passengers.
lb
Manufacturing Weight Empty
204,000
Crew (8 Persons plus Baggage)
1,580
Unusuable Fuel and Oil
1,400
Cooling Water
2,000
Engine Oil
285
Liquid Nitrogen Fuel System
400
Flight supplies, emergency equipment etc
1,945
Operating weight empty (rounded values)
212,000
Payload (Passengers, baggage, freight)
31,000
Operating Zero fuel Weight
243,000
Reserve fuel (for 300 n.m. diversion at cruise altitude plus 30 minutes hold at sea level
25,000
Operating Landing Weight
268,000
Flight Fuel
171.000
Maximum all-up weight
439,000
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KJ_Lesnick
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Does anyone have any diagrams or pictures of models, etc, of the B-2707-100 when they widened the fuselage of the winning design to the point that it could accomodate 2-3-2?
KJ
KJ
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Hi,
the Lockheed L-2000 was displayed here with large fuselage and
small fuselage variants.
the Lockheed L-2000 was displayed here with large fuselage and
small fuselage variants.
sst | 1966 | 0144 | Flight Archive
International, 20 January 1966 95 MACH THREE TECHNOLOGY Lockheed's Supersonic Trans- port: Part 1—The Background; Aerodynamics; Propulsion; Air- frame Structure General arrangement of the Phase 28 Loc
www.flightglobal.com
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KJ_Lesnick
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What was the airplane's length with the shorter fuselage? How did that affect the area-ruling?
BTW: Was that with the 9,026 square-foot wing-area used in the L-2000-2, or the 9,424 square foot wing-area used in the L-2000-7A?
BTW: Was that with the 9,026 square-foot wing-area used in the L-2000-2, or the 9,424 square foot wing-area used in the L-2000-7A?
KJ_Lesnick
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The L-2000-7A's leading-edge droop also required the trailing-edge devices to deflect down (or at least upward at a lesser deflection) no?
Sentinel Chicken
American 71 Heavy, contact departure 126.47
Wonderful document, Skybolt! I've been trying to work on a Boeing 733 print for sometime now, this will come in VERY handy!
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ha ah, finally got it...
PROPOSAL FOR THE DEVELOPMENT OF A COMMERCIAL SUPERSONIC TRANSPORT. VOLUME A-5. AERODYNAMIC REPORT - LOCKHEED-CALIFORNIA CO BURBANK - 15 JAN 1964
Part of the original Phase I proposal. Photo reproductions are abysmal, but line drawings are good. Some of the photos later published in Horizons magazine.
apps.dtic.mil
PROPOSAL FOR THE DEVELOPMENT OF A COMMERCIAL SUPERSONIC TRANSPORT. VOLUME A-5. AERODYNAMIC REPORT - LOCKHEED-CALIFORNIA CO BURBANK - 15 JAN 1964
Part of the original Phase I proposal. Photo reproductions are abysmal, but line drawings are good. Some of the photos later published in Horizons magazine.
Defense Technical Information Center
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KJ_Lesnick
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Cool...
You know I never even heard of the Curtiss Wright TJ70A-4 (one of the engine's to power the Lockheed L-2000 Phase 1). Fascinating
You know I never even heard of the Curtiss Wright TJ70A-4 (one of the engine's to power the Lockheed L-2000 Phase 1). Fascinating
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Actually the Curtiss-Wright was the most advanced of all three engine proposals, but it was deemed not available in time for the original assumed schedule of the SST (first flight in 1971). There was a lot of controversy, with NASA and Air Force insisting in keeping Curtiss-Wright in the game with funds for building a second-generation supersonic transport engine for the 1975 time-frame. But the FAA was racing with Concorde, so Curtiss-Wright was dropped. BTW, the company put significant resources in the design of that engine, but to no avail. The Air Force had already shifted to highly-subsonic low level penetration with incoming and outcoming supersonic legs for their post-XB-70 bombers concepts.
KJ_Lesnick
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When I did web-searches on this one, I don't recall ever seeing an explanation regarding the HSCT's weight, just that the plane's weighed a given amount. I have even asked on other web-forums regarding the following particular questions...
-How come the HSCT designs came out weighing something like 738,000 lbs utilizing a much more advanced composite design and no need for a droopable-nose?
-How is it that the B-2707-300 which originally started out at 635,000 lbs using a tailed double-delta wing, ended up creeping all the way up to 750,000? (I know with the B-2707-100 and B-2707-200 the creep up ended up occurring because of the swing-wings.) It was a fairly simple-design --- how did they screw up so much?
-Would the Lockheed L-2000-7A have ended up drastically heavier than the 590,000 lb estimate like the double-delta B-2707-300?
but I never got a particularly good answer....
Kendra Lesnick
-How come the HSCT designs came out weighing something like 738,000 lbs utilizing a much more advanced composite design and no need for a droopable-nose?
-How is it that the B-2707-300 which originally started out at 635,000 lbs using a tailed double-delta wing, ended up creeping all the way up to 750,000? (I know with the B-2707-100 and B-2707-200 the creep up ended up occurring because of the swing-wings.) It was a fairly simple-design --- how did they screw up so much?
-Would the Lockheed L-2000-7A have ended up drastically heavier than the 590,000 lb estimate like the double-delta B-2707-300?
but I never got a particularly good answer....
Kendra Lesnick
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The give that answer would need some hundreds pages. I give just an hint for the 2707: structural weight of the wing box and pivots. The -200 was abandoned when calculations showed that design convergence would have required a TOW of more than 1.2 million pounds. As for the HSCT, it was combining low and high speed efficiency and amount of fuel the major culprits, not to mention the sonic boom reduction (very slender airframe, with consequent stiffening of the structure) and sound suppression at low altitude. Build a somewhat "green" SST is VERY difficult, expecially if you want to make it economically viable for the airlines.
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What really hit the 1990s HSCT badly was the requirement for Stage 3 noise compliance, which nailed you in several ways. The engine cycle had to be compromised: at Mach 2.4 what you really want is nearly a straight TJ, but that would never make the noise requirement. The chosen solution was a turbofan with a giant, heavy suppressor, which was a weight penalty. Then you have the escalation factor of all that weight hanging off the back of the wing structure.
I believe the reference range was also greater than for the final Boeing SST in the early-1970s program.
I believe the reference range was also greater than for the final Boeing SST in the early-1970s program.
KJ_Lesnick
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Skybolt,
I never knew the weight of the swing-wings would actually get that high! The B-2707-300, though, was a tailed double-delta design... how did it's initial estimate of 635,000 lbs go all the way up to 750,000 lbs? It's range to my knowledge was roughly the same as the earlier SST design specs from what I remember...
Also, while I'm at it... would the L-2000-7A have been able to make it's 590,000 MTOW weight figures (or at least come close)? I'm just wondering because, while the design was fairly simple, and Lockheed had more experience with supersonic designs, and titanium fabrication than most other aircraft manufacturers, during the 1970's, Lockheed tried to develop an SST design again (I'm not sure the exact details, although at least one was to use LH2), that was a tailed SCAT-15F type design that was 286 or 287 feet long, powered by 4 x 80,000 lbf engines (what engine that was I got no idea), and weighed 750,000 lbs! Was this design so heavy because Lockheed realized how much such a design would really weigh, or were they right about the L-2000-7A, but because of the (tailed) cranked-arrow wing design (which was rejected during the L-2000 program) and it's weight-penalties over the (tailless) delta-wing?
Kendra Lesnick
BTW: I have an ENORMOUS amount of information accumulated over the years about the SST program -- I have never found an answer for these questions in the information I possess... internet searches have not provided me with this information either...
I never knew the weight of the swing-wings would actually get that high! The B-2707-300, though, was a tailed double-delta design... how did it's initial estimate of 635,000 lbs go all the way up to 750,000 lbs? It's range to my knowledge was roughly the same as the earlier SST design specs from what I remember...
Also, while I'm at it... would the L-2000-7A have been able to make it's 590,000 MTOW weight figures (or at least come close)? I'm just wondering because, while the design was fairly simple, and Lockheed had more experience with supersonic designs, and titanium fabrication than most other aircraft manufacturers, during the 1970's, Lockheed tried to develop an SST design again (I'm not sure the exact details, although at least one was to use LH2), that was a tailed SCAT-15F type design that was 286 or 287 feet long, powered by 4 x 80,000 lbf engines (what engine that was I got no idea), and weighed 750,000 lbs! Was this design so heavy because Lockheed realized how much such a design would really weigh, or were they right about the L-2000-7A, but because of the (tailed) cranked-arrow wing design (which was rejected during the L-2000 program) and it's weight-penalties over the (tailless) delta-wing?
Kendra Lesnick
BTW: I have an ENORMOUS amount of information accumulated over the years about the SST program -- I have never found an answer for these questions in the information I possess... internet searches have not provided me with this information either...
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Its pretty common for initial weight estimates to be optimistic. Nothing odd or unusual about it.
KJ_Lesnick
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Overscan,
Just out of curiousity, do you think the L-2000-7A's listed weight-figures were as "optimistic"? I'm just curious as...
- The L-2000-7A is, in many ways, simpler (no horizontal-stabilizer, no flaps, and less complicated droops, it may have had a better structural-design).
- Lockheed had more experience with supersonic aircraft, with titanium fabrication, and with environmental-control, and cooling systems (To keep the passengers, pilots and flight-attendants from getting roasted) and would be better able to estimate/guesstimate weight.
Kendra Lesnick
Just out of curiousity, do you think the L-2000-7A's listed weight-figures were as "optimistic"? I'm just curious as...
- The L-2000-7A is, in many ways, simpler (no horizontal-stabilizer, no flaps, and less complicated droops, it may have had a better structural-design).
- Lockheed had more experience with supersonic aircraft, with titanium fabrication, and with environmental-control, and cooling systems (To keep the passengers, pilots and flight-attendants from getting roasted) and would be better able to estimate/guesstimate weight.
Kendra Lesnick
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The 2000-7a and b had other problems as well. As far as noise, it didn't met the already untenable DC-8 and 707 standard. Moreover, its payload at 4000 miles sunk to 30.000 lbs due to ever increasing demand for reserve fuel. The double-delta had problems at subsonic speeds and the airlines were upset by the need to remain aloft circling waiting for landing on increasingly congested airports. Moreover, the climb rate was lower than the Boeing's one, so the noise footprint problem got even worse. Finally, and this became apparent at the Phase III evaluation, the 2000-7 was weight-wise at the limit, i.e. it had no more room for weight increases, any of those would have required a complete redesign.
As a side note: a real complete, technical AND political AND economical history of the US SST program has yet to be written. A good first work is "High Speed Dreams", by Erik M. Conway, John Hopkins University Press, 2005, ISBN 0-8018-8067-X, but it covers ALL officially-backed (NASA primarily) SST efforts in the US, so it must cut short on details. Highly recommended. though.
As a side note: a real complete, technical AND political AND economical history of the US SST program has yet to be written. A good first work is "High Speed Dreams", by Erik M. Conway, John Hopkins University Press, 2005, ISBN 0-8018-8067-X, but it covers ALL officially-backed (NASA primarily) SST efforts in the US, so it must cut short on details. Highly recommended. though.
KJ_Lesnick
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Skybolt said:
The B-2707 did?
What was the initial 4,000 nm estimate without the increased reserve-fuel requirements?
What problems? I thought they got rid of the pitch-up problems by Phase II.
Hard to believe as the B-2707 had a lower T/W ratio. What were their estimated climb-rates anyway?
I guess the L-2000's relatively poor-climb a result of the high-alphas inherant in the delta-wing...
I never knew that it was right at the limit. Was the 590,000 lb figure even remotely accurate as it was?
KJ_Lesnick
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KJ, those were the judgments given by the Phase III selection committee in 1966...
As for noise, the B-2707 produced less noise on the airport area than a 707, due to the higher climb and sink rate (less footprint). The strictly engine exhaust-related noise was more or less the same. As for the other answers, I don't have here my sources, so please wait some hours.
As for noise, the B-2707 produced less noise on the airport area than a 707, due to the higher climb and sink rate (less footprint). The strictly engine exhaust-related noise was more or less the same. As for the other answers, I don't have here my sources, so please wait some hours.
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Yes
Add 25 per cent, and before you ask, with equivalent range fuel reserves (subsonic performance was better) the 2707 scored 60.000.
Less efficient than the 2707, higher fuel consumption
Hard but true: problem was weight
As I said before, a comprehensive technical history of the SST effort still lacks; but the evaluation team, and a further independent team, said exactly this. 590.000 was the final datum in the Lockheed Phase III proposal.
As a final round-up, if someone is really interested in re-evaluating the two contenders I think they can read all the Phase III documents (Boeing's are mostly available online, Lockheed's are in the FAA archive) and make up their minds. A good starting point could be the already cited "High Speed Dream" by Erik M. Conway.
No-one of the contenders (and I include the engines) would have been able to provide the performance requested with the noise level requested AND THE ECONOMICS requested by the airlines. Even not considering the sonic boom. Already in late 1966 the DoD was internally sure that no civil airliner would have been permitted to fly supersonically overland. And FAA suspected the same, even asking Boeing to do research on highly subsonic and no-boom supersonics (Mach 1.2) to fill tha gap.
KJ_Lesnick
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Then who said the B-2707 was 675,000 MTOW? Same people?
KJ_Lesnick
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- 104
Skybolt,
Am I reading this wrong, or did the same people that said the L-2000 would have a 590,000 max weight said that the B-2707 woul d have a 640,000 lb max weight?
Am I reading this wrong, or did the same people that said the L-2000 would have a 590,000 max weight said that the B-2707 woul d have a 640,000 lb max weight?
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