This paper describes the design features of a Douglas Mach 2.4/1.8 Low Sonic Boom High Speed Civil Transport (HSCT) configuration developed for NASA. The configuration is designed to fly over water at Mach 2.4 for highest productivity and economic worth, and fly over land at Mach 1.8 with reduced sonic boom loudness

Morgenstern, John M.

English

NASA Technical Reports Server

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LOW SONIC BOOM DESIGN AND PERFORMANCE OF A MACH 2.4h.8 OVERLAND 
HIGH SPEED CIVIL TRANSPORT 
4 
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John M. Morgenstern 
McDonnell Douglas Corporation, Douglas Aircraft Company 
Long Beach, CA 
INTRODUCTION 
This paper describes the design features of a Douglas Mach 2.4/1.8 Low Sonic Boom High 
Speed Civil Transport (HSCI' )  configuration developed for NASA under government contract 
number NAS 1- 19345. The configuration is designed to fly over water at Mach 2.4 for highest 
productivity and economic worth, and fly over land at Mach 1.8 with reduced sonic boom loudness. 
SONIC BOOM DESIGN GOALS 
Before the design work was undertaken, a study was performed to determine the appropriate 
Mach number and weight for sonic boom design minimization efforts (figure 1). Based on 
preliminary acoustic response studies, a loudness goal of 90 PLdB was chosen. The NASA Langley 
(Christine Darden) SEEB computer code* was used to quantify the maximum weight possible for a 
90 PLdB waveform. A minimum shock waveform was chosen to maximize the weight allowable. A 
minimum weight constraint was added based on a non-low boom baseline configuration weight 
required to meet the mission range and payload desired. As a first order approximation, a 10% 
increase in weight above the baseline was chosen as an upper weight limit. In previous NASA 
studies, Douglas was counseled against operating in the no bow shock regime due to concerns over 
shockless waveforms coalescing into larger front shocks. Further, the equivalent area shapes 
required to eliminate bow shocks tend to allow very little volume for a practical vehicle. From the 
design space shown in figure 1, it was decided to pursue sonic boom minimization at Mach 1.8 with 
a maximum take-off gross weight (MTOGW) of 850,000 pounds. 
In addition to using Mach 1.8 for sonic boom minimization, a payload of 300 passengers and 
r&ge of 5,000-5,500 nautical miles were chosen to match the baseline configuration. Mach 2.4 
supersonic cruise over water was chosen to reduce the risk of the low boom design: if the low boom 
https://ntrs.nasa.gov/search.jsp?R=19920076745 2020-03-24T07:22:36+00:00Z
Sonic Boom Design Space 
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1.4 s 18 20 22 24 
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- Front Shock Yin. Shape 
- bngth = 330 n. 
- P r m  I 0.5 
- Yf/L 0 0.0 . 
kllach 
Figure 1 
configuration was not allowed to fly supersonic overland, it would still be capable of Mach 2.4 
over water; therefore, it retains the baseline's productivity economic advantage over water (compared 
to subsonic aircraft). Further, the long lifting lengths needed for sonic boom minimization are 
conducive to the higher sweeps which are beneficial at Mach 2.4. 
DESIGN AND ANALYSIS METHODS 
Design of the configuration started with a SEEB equivalent area goal based on the chosen MBch 
number and weight. A linear wing-body-tail panel method calculates the configuration's lift in a 
trimmed condition. A far-field wave drag program calculates the volume pressure disturbance of 
the full configuration, wing-body-tails-nacelles. A linear sonic boom propagation method reads in 
axial lift distributions at several angles-of-attack and Mach angle cut volume distributions at several 
roll an les. The lift and volume are converted into an 'f function (as per the Whitham-Walkden theory F ) and propagated to the ground at several roll angles to fonn a sonic boom carpet. 
Numerous configurations, their permutations and various mutations were analyzed. This study 
included at least 3 planforms, several twist and camber combinations, 3 horizontal tails, 4 canards, 2 
vertical tails, 2 nacelles, and more than 50 fuselages. Many improvements were made in the analysis 
methods to handle more complex configurations, improve turn-around time, and improve techniques 
used for numerically calulating slopes and second derivatives as configuration complexity increased. 
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Pmx - 2.2 
ROB - 93.5 
E E L -  102.6 I 
PWiX - 2.5 
ROB- 98.4 
EEL- 104.6 
MACH 1.8 GROUND WAVEFORM MACH 1.8 GROUND WAVEFOFM 
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9. 
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alo.0 Figure 2 a1 0.0 Figure 3 
During the many configuation analyses and recent NASA wind tunnel test analyses: the strong 
effect of nacelles on sonic boom became apparent. Figure 2 shows a comparison of the intended 
minimum shock waveform with the ground waveform for an earlier version of the 2.4/1.8 configura- 
tion. A large pressure spike due to the nacelles "ate up" most of the ramp waveform, creating a 
second shock that doubled the front waveform loudness. A two part strategy was undertaken to deal 
with the nacelle pressure spikes: the location of the nacelles was changed, and the accuracy of 
nacelle effects was increased. 
First, the nacelles were moved aft so their pressure spikes would occur behind the front ramp of 
the waveform. This virtually eliminated the nacelle shock effect on the ground waveform, as shown 
NACELLE PRESSURE CONTOURS FROM NASTD CFD SOLUTION 
- Nacelle Ourline 
Pressure Contours - 
Figure 4 
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m 
f 
MDBOOM PREDICTED 'F' FUNCTION AT MACH 1.8 BEGINNING OF CRUISE 
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in figure 3. However, it necessitated a large change in the configuration. To get the nacelles back 
far enough, a change was made to a canard-wing arrangement (instead of an aft tail as on the 
baseline) and a new planform was developed. This arrangement allowed the airplane balance to be 
achieved with a more aft wing placement and allowed the typical underwing nacelle mounting 
arrangement to be retained. 
Finally, a new method was developed to improve the modelling of nacelles and nacelle 
interference effects. CFD runs were made of a detailed nacelle and diverter installation (developed 
for the NASA Lewis Propulsion Airframe Integration Technology, PAIT, study), as shown in figure 
4. The pressure field from the Euler CFD solution (NASTD run by McDonnell #iircraft; St. Louis, 
MO) is used in the linear 'f function calculation and this helps in modelling interference effects. The 
impact of including nacelle effects on the 'f function calculation is shown in Figure 5. Similar 
trends have been obtained in recent NASA wind-tunnel tests3 
MACH 2.411.8 GENERAL ARRANGEMENT 
PLAN VIEW 
- -  * 
FRONT VIEW 
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SIDE VIEW - m 
Figure 6 I 
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MACH 2.411.8 GROUND UNDERTRACK WAVEFORM 
AT MACH 1.8 BEGINNING OF CRUISE 
1 GOAL--& \ 
P I 
- =ImE' ""Q; 
0.a 0.a Pp .e 
ms 
Ba'b.0 0.1 R r  
BACK HALF 
97.5 PLdB 
: 10.0 
Figure 7 
2.4/1.8 CONFIGURATION GENERAL ARRANGEMENT 
The latest 2.4/1.8 configuration arrangement is shown in figure 6. As mentioned 
previously, it has a canard in front. In addition to acting as a trimming and control surface, the 
canard has a small movable surface on its zero sweep leading edge. This surface is deflected up 
during low boom operation at Mach 1.8 to help create an effective "nose blunmess" without a blunt 
area distribution. In this way, the bluntness drag need not be present during Mach 2.4 over water 
cruise. The wing has a 76/68 degree subsonic swept leading edge and a "gull-wing'' dihedral to 
imppve nacelle clearance. The fuselage meets the 300 passengers in a 3-class arrangement goal. 
2.4/1.8 CONFIGURATION BOOM LEVELS 
Figure 7 shows the beginning of cruise undertrack ground waveform. While the front half of the 
waveform is close to the 90 PLdB goal, the aft shock brings the total to an annoying 98.2 PLdB. . 
The first question is, "Why did the aft shock get so loud?" The loud aft shock appeared as a result of 
adding the nacelle CFD pressures created under the back end of the wing. These pressures cause a 
steep rise in the equivalent area distribution seen as the nacelle pressure spike in the 'f function 
(figure 5). As previously mentioned, the effect of the spike is suppressed by the nacelle placement; 
however, the quick drop off in lift behind the spike causes negative pressure spikes behind the 
nacelles, which could not be eliminated by fuselage shaping without a large drag penalty. A more 
gentle drop off in lift is needed to reduce the aft shock strength. 
The second question might be, "Why has the minimum shock waveform become a delayed 
ramp?" The front shock is a little larger than desired because the large canard at the front nose 
introduces some non-smoothness that is difficult to remove. Also, the design methods have trouble 
keeping the area distributions smooth wherever lifting surfaces begin and end. And finally, the 
shape of the ramp was allowed to vary a bit to smooth the fuselage for reduced wave drag. 
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-15.0 -10.0 -5.0 0.0 5.0 10.0 15.0 
DISTANCE OFF-TRACK (FEET*lOOOO) 40' 
Figure 8 
(original figure unavailable) 
One solution to both problems would be to add an aft horizontal tail in addition to the front 
canard. In an optimum supersonic trim condition, it is generally beneficial to carry some aft tail lift. 
able methods are not currently capable of handling three surface configurations. 
In spite of the large aft shock, it is important to hightlight how far the design has progressed 
toward an acceptable low boom signature. Significant attention was paid to minimizing off-track 
loudness during the design process. Figure 8 shows the total sonic boom carpet loudness for the 
beginning of cruise condition, along with waveforms at several points along the carpet. Note that a 
shaped waveform is retained throughout the carpet. The effect of the nacelle pressure spike is 
controlled at all roll angles. 
In addition to looking at off-track loudness, sonic boom carpets at other than the beginning of . 
cruise must be considered. Figure 9 compares the Mach 1.8 beginning of cruise sonic boom carpet 
with the Mach 1.2 climb carpet (without acceleration effects) and the Mach 1.8 end of cruise carpet. 
The Mach 1.2 climb values are held to roughly the same loudness level, but the carpet is consider- 
- The size of the canard could also be reduced. The problem in using this approach is that the avail- 
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60 
SONIC BOOM CARFET 
MACH 2.4/1.8 
. 110.0 
105.0 
a 
3 :OD.: 
E 
4 
9 
r n  
z 
4 95.0 
90.0 
05.0 
80.0 
-15.0 -10.0 -5.0 0.0 5.0 10.0 75.0 
DISTANCE OFF-TRACK (FEET+lOOOO) *lo' 
Figure 9 
ably more narrow. A low noise, mid-field (non N-wave) waveform at the ground is achieved by 
climbing in a low q (dynamic pressure) ascent tragectory similar to those shown in reference 3. The 
fuel burn penalty for the lower q climb is 1.5 to 2 percent. It is possible that this penalty could be 
reduced by tailoring the canard lift dismbution for Mach 1.2 boom minimization in future studies. 
* 
Of paticular interest is that the Mach 1.8 end of cruise waveforms are almost identical in shape to 
the beginning of cruise waveforms, except they are about 6 PLdB quieter. The ground waveform 
from the end of cruise has the same "aged" shape as the beginning of cruise waveform, because the 
airplane is flying at a constant lift coefficient which means that the same pIJp, ratio exists all along 
the vehicle throughout cruise. The implication is that throughout climb, cruise, and descent the 
"aging" of a low boom ramp waveform will be held at or below its design target, and thereby, 
maintain a shaped waveform (in a non-turbulent atmosphere.) In summary, altitude changes at the 
same Mach number do not cause low boom waveform shapes to "age" differently when using typical 
constant lift coefficient climb, cruise, and descent mgectories. Conversely, Mach number directly 
affects waveform "aging", so that at Mach 2.4 beginning of cruise the low boom 2.4/1.8 
configuration produces a typical N-wave of 104.5 PLdB loudness, undertrack. 
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2.4/1.8 CONFIGURATIONPERFORMANCE 
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Since Mach 2.4 cruise over water is more productive and 70% of the earth is cbvered With water, 
be  2.4/1.8 low boom airplane is likely to spend most of its time cruising at Mach 2.4. Therefore, 
Mach 2.4 cruise efficiency is of paramount importance. Wing leading edge sweeps were chosen 
with this speed in mind and the planform twist and camber distribution are optimized for it. The 
area dismbution is smoothed for Mach 2.4 cruise and simultaneously shaped for low boom at Mach 
1.8. The Mach 2.4 roll averaged area distribution is shown in figure 10. 
. 
MACH 2.411.8 ROLL AVERAGED AREA DISTRIBUTION AT MACH 2.4 
X STATION 
Figure 10 
Table 1 below gives a comparison between the 2.4D.8 low boom airplane and a 2.4B.95 
baseline, both sized at the same MTOGW and flying 25% over land at reduced speed. The Mach 2.4 
cruise perfoxmance of the two configurations is about equal; however, the 2.4/1.8 configuration is 
significantly more efficient at Mach 1.8 over land than the 2.4B.95 baseline is at Mach 0.95 over 
Iand. The main penalty of the 2.4/1.8 configuration is the weight of the large wing, which contrib- 
utes most of the 18 percent increase in operator's empty weight (OEW) relative to the Baseline. This 
performance assessment analysis is preliminary and no changes of the Low Boom geometry were 
allowed during sizing. 
Table 1 
LOW BOOM 
2.411.8 
BAS ELI N E 
2.410.95 
830 000 830 000 
358 000 304 000 
66 89 
62 900 57 900 
5 000 5 660 
---------------- ----------------- 
@ 25% OVERLAND 
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WIND TUNNEL TEST, ND FURTHER PLANS 
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At this point, it is desirable to fabricate and test a wind tunnel model of the 2.4/1.8 low boom 
geometry. It represents many new features that have yet to be validated. Predictions indicate that a 
ramp front waveform should be achieved undertrack and shaped waveforms are achieved at all roll 
angles with ground intercepts. The strong aft shock would not make a difference with the proposed 
wind tunnel test methods, because the signature behind the nacelles cannot be modelled accurately 
due to sting interference and the lack of engine power effects (engine exhaust). Further revisions of 
the 2.4/1.8 configuration will be undertaken to reduce the aft shock and improve performance. In 
the meantime, the CFD predictions of the nacelle effects need to be verified along with the off-track 
prediction methodology. Further, a newly developed method that links full configuration CFD 
solutions to sonic boom predictions4 will be used on the 2.4/1.8 Low Boom design. The CFD is 
better able to account for three dimensional effects, but more complex, low boom wind tunnel test 
data are needed to develop experience and thereby accuracy with the method. 
REFERENCES 
1. Darden, Christine M.: Sonic-Boom Minimization With Nose-Bluntness Relaxation. NASA 
'I"-1348, Jan. 1979. 
2. Whitham, G. B.: The Flow Pattern of a Supersonic Projectile. Communications on Pure and 
Applied Mathematics, 5, 1952, pp. 301-348. 
Walkden, F.: The Shock Pattern of a Wing-Body Combination, Far From the Flight Path. 
Aeronautical Quarterly, TX (2), 1958, pp. 164-194. 
3. Darden, Christine M.; Mack, Robert J.; Needleman, Kathy E.; Baize, Daniel G.; Coen, Peter 
G.; Barger, Raymond L.; Melson, Duane N.; Adams, Mary S.; Shields, Elwood W.; and 
McGraw, Marvin E.: Design and Analysis of Low Boom Concepts at Langley Research 
Center. First Annual High-speed Research Workshop, Vol. III, Session 5,  May 1991. 
4. Page, Juliet A.; Plotkin, Kenneth J.: An Efficient Method for Incorporating Computational 
Fluid Dynamics Into Sonic Boom Prediction. AIAA 91-3275, Sept. 1991. 
ACKNOWLEDGEMENT 
The author would like to acknowledge the help of the Advanced Design Team of Douglas 
Aircraft. 
63 


Low sonic boom design and performance of a Mach 2.4/1.8 overland high speed civil transport

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