N AS A / TM- 1999-209535

A Rapid Empirical Method for Estimating

the Gross Takeoff Weight of a High Speed

Civil Transport

Robert J. Mack

Langley Research Center, Hampton, Virginia

December 1999

https://ntrs.nasa.gov/search.jsp?R=20000023189 2018-05-14T14:10:01+00:00Z
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NASA/TM-1999-209535

A Rapid Empirical Method for Estimating

the Gross Takeoff Weight of a High Speed

Civil Transport

Robert J. Mack

Langley Research Center, Hampton, Virginia

National Aeronautics and

Space Administration

Langley Research Center

Hampton, Virginia 23681-2199

December 1999
Available from:

NASA Center for AeroSpace Information (CASI)

7121 Standard Drive

Hanover, MD 21076-1320

(301) 621-0390

National Technical Information Service (NTIS)

5285 Port Royal Road

Springfield, VA 22161-2171

(703) 605-6000
SUMMARY

During the cruise segment of the flight mission, aircraft flying at supersonic speeds generate

sonic booms that are usually maximum at the beginning of cruise. The pressure signature with the

shocks causing these perceived booms can be predicted if the aircraft's geometry, Mach number,

altitude, angle of attack, and cruise weight are known. Most methods for estimating aircraft

weight, especially beginning-cruise weight, are empirical and based on least-square-fit equations

that best represent a body of component weight data. The empirical method discussed in this

report used simplified weight equations based on a study of performance and weight data from

conceptual and real transport aircraft. Like other weight-estimation methods, weights were

determined at several points in the mission. While these additional weights were found to be

useful, it is the determination of beginning-cruise weight that is most important for the prediction

of the aircraft's sonic-boom characteristics.

INTRODUCTION

Weight estimation is one of the most essential, yet one of the most difficult tasks required

during the design of an HSCT concept. It is of particular importance when designing a low-sonicboom concept because its beginning-cruise weight must be known for both aerodynamic

performance and sonic-boom characteristics analyses. The results of these analyses often lead to

modifications which impact the configuration weight during the entire mission including that at

the beginning of cruise. A simple yet effective method for estimating beginning-cruise weight

would be a help during this process even though the weight would be back-of-the-envelope

accurate. As the configuration took shape, more exact methods could supply weight estimates so

that a design solution which satisfied both sonic-boom and mission requirements would quickly

converge.

The method described in this paper for estimating mission weights is based on weights from

similar real aircraft and conceptual HSCT designs that fly similar missions. The weights in this
data base provide estimates that are consistent with the accuracy of the initial input data. General

design perturbations for reducing the sonic boom can be introduced and the trends in weight

change, due to these perturbations, can be estimated so that sonic boom characteristics and

mission performance data can be determined simultaneously.

a

b

B

C

e

kc

kR

In

L/D

M

n

R

RClimb

RDesc

S

W

WCR,e

SYMBOLS

speed of sound at altitude, ft/sec

wing span, ft

Brequet range factor, (M a / c) (L / D), nmi, equation (11)

mean specific fuel consumption, Ibfuel/lbthrust-hr

basis of natural logarithms, 2.71828...

structural aspect ratio factor, defined in figure 2

ratio of WF, Climb/WGT O

ratio of WERes/WGT O

natural logarithm

mean lift to drag ratio during cruise

Mach number

number of passengers

total range, nmi

range in climb and acceleration, nmi

range in cruise, nmi

range to descend, nmi

wing area, ft2

weight, Ib

weight at end of high-speed cruise flight, lb

2
WCR,i

Wcrew

WDesc

WE

WF, Climb

WF, CR

WF, Res

WGTO

Wpass

Z

A0.75

weight at beginning of high-speed cruise flight, lb

weight of pilot, first officer, and flight attendants, lb

weight of fuel required to descend and land, lb

empty weight, lb

weight of fuel used to takeoff, climb, and accelerate to cruise Mach number, lb

weight of fuel used during high-speed cruise, WCR,i - WCR,e, lb; equation (10)

weight of reserve fuel, lb

gross takeoff weight, lb

weight of passengers and baggage, lb

exponent of e in the range equation

sweep angle of the line between the 75 percent stations of the root and tip chords

DISCUSSION

The weight of a conceptual aircraft depends on its mission. In this paper, a simplified mission

consisting of takeoff and climb, cruise, descend and land, figure 1, is used to obtain an estimate of

start-of-cruise weight. However, to obtain this weight estimate, it is also necessary to calculate

end-of-cruise weight, landing weight, empty weight, and gross takeoff weight.

Altitude

Cruise

a_aeOff, climb,

ccelerate

Figure 1. Mission profile.

Mission Range

The cruise segment of the mission is flown at a constant supersonic Mach number, and begins

immediately after completion of the climb and acceleration segment of the mission. The segment

3
of themissiondescribing a diversion to an alternate airport is not shown, but the required fuel is

included in the reserves.

The gross takeoff weight of a conceptual HSCT can be written as

WGT O = Wpass + Wcrew + WF, Climb + WF, Desc + WF, CR + WF, Re s + W E. (1)

Each of these weight items is discussed in the following paragraphs.

Passenger Weight, Wpass. The weight of the total number of passengers and their baggage is

usually estimated from an average weight of 210 lb per passenger. However, most of the desired

routes will be intercontinental so passengers will probably be given an extra 20 pounds of

luggage. Using the higher luggage allowance,

Wpass = 230 n. (2)

Crew Weight, Wcrew. The weight of the crew, Wcrew, is estimated from 450 lb of pilot and first

officer weight added to the weight of 150 lb per flight attendant. For an HSCT aircraft, one flight

attendant for every 30 passengers is considered average for commercial service. So

WCrew = 450 + 5 n. (3)

Fuel Used During Climb, WClim h. Data from references 1 to 5 indicate that from 6 to 12 percent

of the gross takeoff weight, WGTO, is used to take off, climb, and accelerate to start of cruise. If

high-technology engines are used on the concept, 6 to 8 percent is a reasonable fraction. However,

technology based on long, reliable engine life indicates that 10 to 12 percent is appropriate. Low

aspect ratio wings are ideal for supersonic cruise, but poor for the subsonic segments of the

mission. Using a general approach

WF, Climb = k C WGTO (4)

where k c can be in the 0.06 to 0.15 range. During this flight segment, an average of from 300 to

400 nautical miles, nmi, will be travelled. This estimate of climb-and-acceleration range, like the

estimate of climb-and-acceleration fuel weight, was obtained from references 1 to 5.

An average climb angle of 2 to 3 degrees was obtained from a study of several flight profiles.

Mach number and noise constraints imposed dynamic pressure and throttle limits on the climb

and accelerate phase of the mission, constraining the average climb angle further. Using

conservative start-of-cruise altitudes, a range increment of

RClim b = 300 to 400 nmi (5)

was found to be a reasonable initial estimate for many conceptual low-boom concepts. Thrust-toweight ratio needed to be in the 0.25 to 0.30 range for this empirical estimate to be reasonable,

especially if a balanced field length of 10,000 to 12,000 feet is required. When the solution for

WGT O is calculated, the engine size can also be estimated from this range of assumed thrust-toweight ratios.

Fuel Used During Descent, WDesc. The body of conceptual HSCT mission weight data showed

that for configurations carrying from 250 to 300 passengers over a mission range of about 5000

nmi, about 5000 to 7000 pounds of fuel, WDesc, are needed to descend and land, i.e.,

4
WEDes c = 5000 tO 7000 lb. (6)

Like many of the other features of the method, this is an oversimplification, but it was found

repeatedly in the fuel scheduling for aircraft of 600,000 to 800,000 lb gross takeoff weight, and

seemed to be a reasonable first approximation for the first round of weight estimation. However,

the method of reference 2 estimated that less than 0.010 WGT O of fuel was used during descent,

and in reference 6, 0.015 WCR,e was the estimate of fuel used during descent. As a first estimate,

equation (6) was used in the method when applied to low-boom concepts.

The flight angle used during the descent segment is usually higher than during the climb-andacceleration segment. Since the aircraft climbs slowly during cruise to maintain aerodynamic

efficiency, it ends the cruise segment higher than when it started. So the descent angle must be

steeper to reduce descent and landing time to a minimum. For this, a descent range of about

RDe s = 200 to 300 nmi (7)

was used as an initial estimate for a typical low-boom concept. The lower value, 200 nmi, gives

an average descent angle of about 3.5 degrees which is comfortably close to a final glide path

angle of 3 degrees often used on landing approaches.

Cruise Fuel Weight, WC_RR. The cruise range was found from

RCR = R - RClim b - RDesc

or, with the estimates of RClirnb and RDesc already mentioned in equations (5) and (7)

RCR = R - (500 to 700) nmi (8)

was used to obtain an estimate of fuel consumed during the cruise segment of the mission. This

mission cruise range is used to calculate beginning-cruise and end-of-cruise weights, WCR,i and

WCR,e respectively, from the Breguet equation

RCR = (M a / c)(L / D) In (WcR, i / WCR,e). (9)

where In denotes the logarithm to base e rather than the logarithm to base 10. The beginningcruise weight has been specified by equation (4), so equation (9) provides the end-of-cruise

weight. Then, the cruise-fuel weight can be estimated from

WF, CR = WcR,i - WCR,e. (10)

Defining B, the Breguet factor, as

B=(Ma/c) (L/D) (11)

where the variables on the right-hand side are performance parameters already determined. The

end-of-cruise weight, WCR,e, carl be determined from

WCR,e = WCR,i (e z) (12)

where the constant, Z, is

Z = - RCR / B. (13)

5
Since the weight, WCR, i, is defined as the gross takeoff weight, WGT O, minus the weight of fuel

required to takeoff, climb, and accelerate to cruise Mach number, WF, Climb, it can be written as

WCR,i - WGT O - WF, Climb = (1.0 - kc) WGT O. (14)

Equations (10), (12), and (14) express WF, CR is a function OfWGT O because both initial and endof-cruise weights, WCR,i and WCR,e, can be expressed as functions of WGT O and the known

range, cruise Mach number, and mean engine-fuel consumption parameters

Reserve Fuel Weight, WRes. A review of mission performance weights showed that the reserve

fuel weight was usually a constant fraction of the gross takeoff weight, i.e.

WF, Re s = kR WGTO. (15)

In the application of this method to a low-boom concept, kR = 0.06 is used to obtain an initial

estimate of WGT O so that WCR,i can be calculated. If a more reserve fuel is desired, values of kR

in the range of 0.06 to 0.065 could be used to determine WGT O and WCR,i.

Empty Weight. W__ The final weight increment needed to complete the gross take-off weight

equation is the empty weight of the aircraft. Several related factors must be considered when

estimating this increment. The first factor is materials technology, i.e. conventional metals versus

composites. A second factor is engine technology where combustion temperatures and cruise

Mach number dictates the choice of metals, bypass ratios, inlet design, and thrust-to weight ratios.

A third factor is noise suppression technology. Noise generated by exhaust gases strongly impacts

nozzle weight and takeoff patterns which, in turn, impacts fuel burned and gross takeoff weight. A

fourth factor is wing planform shape and the structural aspect ratio. This last factor influences and

is influenced by the previous three. In figure 2, a plot of (WGTo/WE) versus a structural aspect

ratio factor, FA, is presented using data from a number of real and conceptual configurations. The

data base for this curve is presented in Appendices A and B.

B

WGTO

WE

h

2 -

1 -

0

0

o o ®

o o to

o o

//_-7_ _ Line joining 0.75 r_ --

//_// and tip chords

_/ "A" O'78 V L b2

FA=

S COS2._. 0.75

I I I I t I I I I

1 2 3 4 5 6 7 8 9

FA

Figure 2. Correlation between (WGTo/WE) and a structural aspect ratio factor, FA.
With only a back-of-the-envelope sketch and some preliminary performance calculations

available, this ratio of (WGTo/WE) can be used to obtain a first estimate of WGT O. Only then can

a beginning cruise weight be obtained for an initial sonic boom prediction. Note that commercial

aircraft like the Boeing 747, the B AC VC-10, the Tu-144, and the Concorde, have (WGTo/W E)

ratios that vary from 2.2 to 2.7. This is a modest range considering that the first two are subsoniccruise, moderate aspect ratio wing, low fuel consumption rate aircraft, and the last two are

supersonic-cruise, low aspect ratio wing, high fuel consumption rate aircraft, with mission ranges

that varied from 3500 to about 5200 nautical miles. For one factor to reflect differences in cruise

Mach number and mission ranges indicates that several design details have been combined to

simplify a complex weight-estimation procedure. Contrast the (WGTo/W E) ratios and mission

ranges of these four aircraft with the (WGTo/W E) ratios and mission ranges of the Boeing 874

concept, the LB-16 concept, and the baseline Mach 2.4 concept, the highest three points on the

chart. The first four aircraft are or were used commercially to carry passengers, while the last

three were conceptual aircraft designed with highly-refined future technology.

Obviously, the curve of (WGTo/W E) vs. FA in figure 2 is not the only possible interpretation or

point-weighing of the data. For this report, the curve drawn was used as the upper limit of a

narrow band with heavy emphasis on the data from the Concorde and Tu-144 supersonic-cruise

aircraft, and to a lesser extent, on the data from the Boeing 747 and VC-10 subsonic-cruise

aircraft. The effect of this weight-ratio data and the projected weight curve on the predicted lowsonic-boom characteristics of conceptual aircraft will be demonstrated in the Application section.

With the aid of the somewhat-optimistic weight-ratio curve shown in figure 2, the empty

weight

WE = WGTO / (WGTofWE) (16)

Can be estimated and used to obtain an estimate of the gross take-off weight, WGT O, and the

beginning-cruise weight, WCR,i. Reference 7 suggested that (WGTo/WE) could be expressed as

an exponential function of WGT O for various types of aircraft. This approach would be useful if

the real and conceptual aircraft data in figure 2 were not available. Other weight-ratio data bases

of similar concepts and aircraft exist, but for this simplified method and its application to sonic

boom prediction, the data which formed the curve fairing in figure 2 are used to obtain an estimate

of WCR,i.

Combining equations (1) through (16) results in

WGT O = (Wpass + Wcrew + WF, Desc ) / ((1.0 - k c) e -RIB - kR - 1.0 / (WGTo/WE)). (17)

This gross take-off weight estimate is obtained without the need for iteration because several

materials technology and the overall weight influences of engine performance have been

combined in the variable, (WGTo/WE). Even with these simplifications, the data curve in figure 2

is not linear although the second-order effects are small.

Since fuel weight is a major item for a long-range HSCT concept, the simplifications

incorporated in the (WGTo]W E) ratio require an accurate estimate of the average specific fuel

consumption, c, during cruise. It should be noted that several engine-technology related factors

kc, kR, and B, not just c, affect the (WGTo/W E) ratio and the beginning-cruise weight. Misleading

results will certainly be obtained if overzealous optimism is given free rein. The applications that

follow demonstrate that this method can provide useful first-estimate results of WCR,i when it is
employedjudiciously.While theestimatedbeginning-cruiseweightwill not beexact,it canbe

usedto obtainthenecessarywing lift, angleof attack,effectivelength,andlift equivalentarea

distributionfor calculatingthepressuresignatureatthebeginning-cruisepoint.Ideally,thisvalue

of WcR,i will be somewhat on the high side. When better weight estimates are made with moreexact methods, WCR,i and the predicted sonic boom will usually be found to be lower than the

initial predictions. This is preferable to the reverse situation since it is easier to reach a converged

solution for the aircraft's weight and sonic boom when the weight is being reduced for an aircraft

of a given size.

APPLICATION

A hypothetical HSCT concept from reference 9 was used to provide geometry and mission

data for the application of equations (1) to (17) and the use of the (WGTo/WE) vs. FA curve in

figure 2. The concept's design had advanced just beyond the back-of-the-envelope stage, and

consisted of a wire-frame three-view drawing, figure 3.

Figure 3. Three-view drawing of the HSCT-11E concept from reference 9.

A fuselage, sized to provide a pilots compartment, seating for 300 passengers seated 5-abreast,

and room for aft-fuselage fuel, had an overall length of about 300 feet. The wing planform had a

span of about 138 feet and an area of about 10,300 square feet for an aspect ratio of about 1.84. A

canard surface near the nose furnished lifting moment for controlling attitude and rotation during

takeoff and landing, but did not provide lift during the cruise segment of the mission. The

structural aspect ratio factor, FA, computed from the sweep angle of a line between the 75 percent

stations of the root and tip chords, was about 3.5, and provided a (WGTo/WE) value of 2.65. Due

to the concept's wing area which was about 9 percent larger than found on many previous

baseline configurations, and the unusual planform shape, the (WGTo/WE) factor was reduced to

2.5 for this application of the weight-calculation method. Engine specific fuel consumption data

was known from generic engine-performance data, but engine size, weight, and thrust would not

be known until WGT 0 had been estimated. Preliminary lift, drag, and pitching moment data had

been calculated from geometry, but the (L / D) was still only an estimate since it included crude

$
estimatesof skinfriction androughnessdragcoefficients.Inputdatafor theinitial estimateof the

concept'smissionweightswereasfollows:

M=2.4

R = 5000.0nmi

L/D=9.2

c = 1.35 lb/lb/hr

n=300

FA= 3.5

(WGTo/WE) = 2.5 (adjusted for planform area and shape)

k C = 0.11

kR = 0.062

RClim b = 400 nmi

RDesc = 300 nmi

WF, Desc - 6000.0 lb

The resulting solution for WGT O, W E, and WCR, i was:

WGT O = 763,708.3 lb

W E = 305,483.3 lb

WcR,i = 679,700.4 lb

which is very close to, but not exactly the same as, the weight estimates given in reference 9

which was obtained by a more complicated process.

From this value of WCR,i, a cruise Mach number of 2.4, an effective length of 300 ft, a

beginning-cruise altitude of 58,000 ft, and a "fiat-top" signature shape, the methods described in

references 10 and 11, subject to the cautions mentioned in reference 12, were used to calculate a

low-boom equivalent-area curve. The equivalent areas represented by this curve includes both

volume and lift contributions. Near the nose volume is the dominant contributor, but toward the

tail of the concept, the lift slowly becomes the dominant contributor. On an HSCT concept, the

maximum areas from lift can be from three to five times larger than the maximum areas from

volume. At the tail of the concept, from 90 to 95 percent of the maximum equivalent area at the

tail comes from the lift; the other 5 to 10 percent is due to the nacelles, boundary-layer wake, and

engine plumes.

These equivalent areas, in turn, were used to predict a base-line low-boom ground-level noseshock strength, Ap, of:

Ap (Base-Line) = 1.05 psf

which became the shock strength for further comparisons. Tail-shock strengths are affected by the

nacelle placement which was only tentative at this time. So, only nose-shock strengths are used to

access the low-boom potentials since from 75 to 80 percent of the effective length must be

tailored to control the nose shock.

The simplified weight-estimation method permits easy, quick studies to be made of the effects

of improvements in aerodynamics, engine performance, and materials technology on sonic-boom

characteristics. If the value of any constant in B, the Breguet factor in equation (11), was changed

such that B was improved by the ratio (2.6/2.5), an increase of 4 percent, the calculated solution

for WGT O, W E, and WCR,i was:

9
WGTO= 694,697.8lb

WE= 277,281.1lb

WCR,i= 618,879.2lb

Thematrixof low-boominput variablesthatprovidedthefirstestimateof ground-levelshock

strengthwereused,changingonly WCR,i, tO get a similar estimate of sonic boom for this second

beginning-cruise weight. The decrease of about 61,400 lb in WCR,i reduced the nose-shock

strength from Ap = 1.05 psf to:

Ap (Improved Breguet Factor) = 0.98 psf.

Now, instead of an improvement in aerodynamics or engine efficiency, a "break-through" in

materials technology was assumed. This improvement was introduced as an increase in the

0NGTO]WE) ratio from 2.5 to 2.6, the same amount of improvement used in the previous (L / D)

ratio calculation to increase the Breguet factor, B. The solution for WGT O, W E, and WCR,i for this

improvement in materials technology was:

WGT O = 662,545.4 lb

W E = 254,825.2 lb

WCR,i = 589,665.4 lb

This new value of WCR,i, a decrease of about 90,000 lb in WCR,i from that of the initial concept,

reduced the nose-shock strength from Ap = 1.05 psf to:

Ap (Improved Structural Efficiency) = 0.95 psf.

If only one improvement were realizable, the reduction in sonic boom would be an

encouragement but not a breakthrough. Should there be a 4 percent improvement in both factors

simultaneously, the reduction in weights would be

WGT O = 609,977.5 lb

W E = 234,606.7 lb

WCR,i = 542,880.0 lb

and nose-shock strengths would drop from Ap = 1.05 psf to:

Ap ((Combined Improvements) = 0.89 psf

all due to the 136,800 lb reduction in WCR, i. The inescapable conclusion is that for aircraft

beginning-cruise weights in the 600,000 to 700,000 lb range with an overall length of about 300

ft, weight reductions of a 140,000 lb result in overpressure nose-shock strength reductions of only

about 0.15 psf. This is not encouraging for aircraft designers who envision supersonic-cruise

concepts that generate only 0.5 psf on the ground while carrying 300 passengers at Mach numbers

of 2.4. However, 1.0 psf overpressure shock is a considerable reduction from 3.0 psf nose-shock

strength generated by a concept configured for maximum aerodynamic and structural efficiency

which is cruising at a similar altitude with the same number of passengers and at the same Mach

number. Should more compromises be required to meet all mission constraints, the shock

strengths can be permitted to drift upward. This is a more feasible strategy than try to force shock

strengths downward through modest and limited geometric tailoring of a heavy concept.

IO
However,thesinglebestreductionin beginning-cruiseweightis obtainedby improvementsin

materialstechnology.Fromapracticalviewpoint,only smallimprovementsin materials

technology,enginetechnology,andaerodynamicefficiencycanbeexpectedwith state-of-the-art

technology.So,beforetheaircraftdesignerinvestmuchtime in extensivesonic-boomanalyses,

thebeginning-cruiseweighttrendsmustbeknownevenif is only to first-orderaccuracy.

Theseexamplesdemonstratetheextentof thechangesin low-boomperformancefrom

changesin beginning-cruiseweightdueto improvementsin materials,enginetechnology,and/or

aerodynamicefficiency.Theyillustratehow suchstudiescaneasilyandquickly bedonewith this

simplifiedweight-estimationmethod.Thus,practicalengineeringlimits to thecruise(L / D) ratio,

engine performance, and materials technology could be explored on a particular configuration so

that their effects on sonic boom and mission range can be credibly predicted before concept

features were frozen in place during the preliminary-design process.

High values (2.75 or higher) of the (WGTo]W E) ratio were clearly out of place in a

preliminary design of a near-term, modest-technology HSCT, but they could employed when it

was desired to predict the results and benefits of potential long-term materials and manufacturing

methods research breakthroughs. For example, they could be used to show the level of technology

required to achieve sonic-boom levels of 0.75 psf and lower, or set the predicted aircraft

performance levels required to achieve the optimistic results set forth in reference 8.

This empirical weight-estimation method is not meant to replace more sophisticated methods

for calculating conceptual aircraft weights. However, these more sophisticated methods require

much more time and effort to supply the necessary input information for calculating sonic-boom

characteristics. By "weeding out" inefficient design features, unnecessary complexity, and

undersized engines at an early stage, the preliminary design can be advanced to the stage where

the extra time and effort needed to use the advanced codes are worthwhile. Then, the increased

accuracy in WGT O, W E, and WCR,i estimates can be used to make refinements in structural

design, engine selection, fuel tank location, landing gear, etc. so that a sonic-boom evaluation can

be made on a concept capable of meeting mission requirements.

CONCLUDING REMARKS

The application of equations (1) to (17) and the curve of (WGTo]W E) vs. FA , a structural

aspect parameter, can provide an initial estimate of the beginning-cruise weight, as well as other

mission weights of a conceptual HSCT geometrically tailored for low sonic boom. From the

estimated beginning-cruise weight, the cruise Mach number, the cruise altitude, and the angle of

attack, a low-boom equivalent area curve can be calculated for the concept. With the calculated

equivalent area curve, the corresponding ground-level pressure signature can be predicted. By

making incremental changes to selected component lengths, shapes, location, and weights, to

mission segment range, and to fuel consumption rates during cruise segments of the overall

mission, trends in weights and their effects mission performance, and predicted ground-level

sonic boom can also be estimated.

11
REFERENCES

1.Fenbert,JamesW.; Ozoroski,Lori E; Geiselhart,Karl A.; Shields,ElwoodW.; and

McElrov,MarcusO.: Concept Development Of A Mach 2.4 High-Speed Civil Transport.

NASA/TP- 1999-209694, 1999.

2. Fettermen, David E.: Preliminary Sizing And Performance Of Aircraft. NASA TM 86357,

July 1985.

3. Babar, Hal T., Jr.; and Swanson, E. E.: Advanced Supersonic Technology Concept AST-IO0

Characteristics Developed In A Baseline-Update Study. NASA TM X-72815, January 1976.

4. Robins, A. Warner; Dollyhigh, Samuel M.; Beissner, Fred L., Jr.; Geiselhart, Karl;

Martin, Glenn L.; Shields, E. W.; Swanson, E. E.; Coen, Peter G.; and Morris, Shelby J., Jr.:

Concept Development Of A Mach 3.0 High-Speed Civil Transport. NASA TM 4058,

September 1988.

5. Haglund, George T.: High Speed Civil Transport Study Of Two Mach 1.7 Low Sonic Boom

Designs. NASA CR 191138, September 1993.

6. Carlson, Harry W.; Barger, Raymond L.; and Mack, Robert J.: Application Of Sonic-Boom

Minimization Concepts In Supersonic Transport Design. NASA TN D-7218, June 1973.

7. Raymer, Daniel E: Aircraft Design: A Conceptual Approach. 2nd Edition, AIAA Education

Series

8. Driver, Cornelius: What Gross Weight And Range For An Advanced HSCT? High-Speed

Research: Sonic Boom, Volume II, NASA Conference Publication 3173, 1992.

9. Mack, Robert J.: Boom Softening And Nacelle Integration On An Arrow-Wing High-Speed

Civil Transport Concept. 1995 NASA High-Speed Research Program Sonic Boom Workshop,

Volume H-Configuration, Design, Analysis, and Testing, NASA/CP-1999-209520, 1999.

10. Seebass, R.; and George, A. R.: Sonic-Boom Minimization. Journal of file Acoustical

Society of America, Vol. 51, No. 2, Pt. 3, February 1972, pp. 686 - 694.

11. Mack, Robert J.: Additional F-Functions Useful For Preliminary Design Of

Shaped-Signature, Low-Boom, Supersonic-Cruise Aircraft. High-Speed Research: 1994

Sonic Boom Workshop. Configuration, Design, Analysis, and Testing, NASA/CP-1999-209699,

1999.

12. Mack, Robert J.: Some Considerations On The Integration Of Engine Nacelles Into

Low-Boom Aircraft Concepts. High-Speed Research: Sonic Boom, Volume II, NASA

Conference Publication 3173, 1992.

12
APPENDIX A.

Weights, Planform Parameters, And Ranges For A Sample Of HSCT Concepts And Long-Range

Aircraft.

HSCT Concents

Configuration M WGTQ. lb W__E,lb WGT/O_/.__WE

Baseline Mach 2.4 2.4 614,302 191,474 3.21

Baseline Mach 3.0 3.0 709,999 263,658 2.69

LB- 16 1.6/2.0 646,356 203,310 3.18

HSCT- 10B 1.8/2.4 662,000 245,000 2.69

Boeing 874 0.9/2.4 666,857 206,527 3.23

Boeing 910 1.7/2.4 823,641 295,076 2.79

Boeing 911 1.7/2.4 778,685 273,807 2.84

Boeing 935 1.7/2.4 731,600 319,300 2.29

HSCT - 8A 1.6 690,000 255,000 2.71

AST 100 2.7 718,000 312,299 2.30

McD - AST 2.2 750,000 310,313 2.42

McD Low Boom 1.8/2.4 830,000 299,200 2.77

W_W_GTo/S Range, nmi

3.7 67.5 6500

3.5 58.3 6500

5.6 69.8 6500

5.2 63.3 5000

2.3 105.7 5000

3.2 82.7 5000

5.3 77.8 5000

6.0 81.3 5000

5.0 57.1 5000

3.8 72.0 4000

2.65 75.0 4400

4.0 66 5000

Commercial Aircraft

Supersonic-Cruise

Concorde 2.0 385,000 141,000

Tu- 144 2.2 396,830 157,500

2.73 2.1 99.8 3800

2.52 2.2 72.7 3500

Subsonic-Cruise

Boeing 747 - 200B 0.85 823,000 371,700

BAC VC - 10 0.85 321,995 146,979

2.21 8.2 150 5190

2.18 8.7 107.1 5040

13
AppendixB. PlanformsOf TheAircraft AndTheConceptsListedIn AppendixA.

,- ,,,,_r

B - 874

B -911

B - 910

Baseline Mach 2.4

S

Baseline Mach 3.0

B - 747

14

°

McD 1.8/2.4

AST- 100

LB - 16

HSCT- 8A

LB - 18 (HSCT- 10B)

#

VC - 10
AppendixB. Concluded.

Concorde

-<=7 ---:_=- jJ

Tu - 144

MCDONNELL DOUGLAS AST B - 935

15
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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

December 1999 Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

A Rapid Empirical Method for Estimating the Gross Takeoff Weight of a

High Speed Civil Transport WU 537-07-21

6. AUTHOR(S)

Robert J. Mack

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Langley Research Center

Hampton, VA 23681-2199

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

8. PERFORMING ORGANIZATION

REPORT NUMBER

L-17743

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NAS A/TM- 1999-209535

11. SUPPLEMENTARY NOTES

12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE

Unclassified-Unlimited

Subject Category 05 Distribution: Nonstandard

Availability: NASA CASI (301) 621-0390

13. ABSTRACT (l_ximum 200 words)

During the cruise segment of the flight mission, aircraft flying at supersonic speeds generate sonic booms that are

usually maximum at the beginning of cruise. The pressure signature with the shocks causing these perceived booms

can be predicted if the aircraft's geometry, Mach number, altitude, angle of attack, and cruise weight are known.

Most methods for estimating aircraft weight, especially beginning-cruise weight, are empirical and based on leastsquare-fit equations that best represent a body of component weight data. The empirical method discussed in this

report used simplified weight equations based on a study of performance and weight data from conceptual and real

transport aircraft. Like other weight-estimation methods, weights were determined at several points in the mission.

While these additional weights were found to be useful, it is the determination of beginning-cruise weight that is

most important for the prediction of the aircraft's sonic-boom characteristics.

14. SUBJECT TERMS

Sonic boom; High-Speed Civil Transport; Supersonic aircraft

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