Aerodynamic design assessment of Strato 2C ... - UFSC Aerodesign

Aerospace Science and Technology 6 (2002) 43–51
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Aerodynamic design assessment of Strato 2C and its potential for
unmanned high altitude airborne platforms
Bewertung der aerodynamischen Auslegung von Strato 2C und dessen
Potential für hochfliegende unbemannte Fluggeräte
Dirk Schawe ∗ , Claas-Hinrik Rohardt, Georg Wichmann
DLR – German Aerospace Center, Institute of Aerodynamics and Flow Technology, Lilienthalplatz 7, 38108 Braunschweig, Germany
Received 2 August 2001; revised and accepted 15 October 2001
Abstract
Currently, there is a large interest worldwide in the development of High Altitude Long Endurance (HALE) Unmanned Aerial Vehicles
(UAVs) for a number of civil and military missions, such as routine weather reconnaissance, surface reconnaissance (forest fires, etc.),
earth observation, border patrol and monitoring, fisheries and wildlife refuge management, chemical and biological agent detection, law
enforcement, disaster assistance and monitoring, telecommunications relay, movie production, agricultural surveying and control, and
provision of targeting information. Passenger and transport airplanes operate at cruising altitudes of maximum 12 000 m where the density is
about 25% compared to sea level. HALE-UVAs are foreseen to operate in the stratosphere at altitudes of 24 000 m, twice as high, where the
density drops to about 3.6% of the sea level value influencing the lift of the aircraft strongly. The environmental conditions in such altitudes
pose strong requirements for the aerodynamic layout and the power plant of an aircraft. In Europe Strato 2C – a manned civil research aircraft
– was until nowadays the only aircraft which reached altitudes above 18 000 m. In this paper Strato 2C’s aerodynamic design and propulsion
layout will be presented and critically reviewed for its suitability for these high altitudes. © 2002 Éditions scientifiques et médicales Elsevier
SAS. All rights reserved.
Zusammenfassung
Derzeit gibt es weltweit ein grosses Interesse für die Entwicklung unbemannter Fluggeräte für zivile, als auch militärische Missionen,
die in grossen Höhen und über einen langen Zeitraum operieren sollen (HALE-UAVs – High Altitude Long Endurance Unmanned Aerial
Vehicles). Angedachte Missionen sind Wetterbeobachtung, Umweltüberwachung (Waldbrände, etc.), Erderkundung, Überwachung und
Schutz der Staatsgrenzen, Kontrolle von Fisch- und Tierschutzgebieten, Nachweis chemischer und biologischer Stoffe in der Atmosphäre,
Strafverfolgung, Katastrophenschutz, Telekommunikation, Filmproduktion, Kontrolle und Überwachung der Landwirtschaft und militärische
Aufklärung und Zielerfassung. Passagier- und Transportflugzeuge operieren in Reiseflughöhen von maximal 12 000 m. In diesen Höhen
hat sich die Luftdichte auf ungefähr 25% des Druckes in Meereshöhe reduziert. Die hochfliegenden unbemannten Fluggeräte sollen in
der Stratosphäre bis in Höhen von 24 000 m betrieben werden. Hier sinkt die Luftdichte auf 3.6% des Druckes in Meereshöhe, wodurch
der Auftrieb des Luftfahrzeuges stark beeinflusst wird. Die Umweltbedingungen in solchen Höhen beeinflussen damit sehr stark die
aerodynamische Auslegung des Flugkörpers und die Auswahl und Auslegung des Antriebes. Bis zum heutigen Tag war das zivile bemannte
Forschungsflugzeug Strato 2C in Europa das einzige Fluggerät, das Flughöhen oberhalb 18 000 m erreichte. In diesem Artikel werden die
aerodynamische und antriebsseitige Auslegung von Strato 2C vorgestellt and kritisch auf ihre Eignung für grosse Flughöhen überprüft.
© 2002 Éditions scientifiques et médicales Elsevier SAS. All rights reserved.
Keywords: Airfoil; Aerodynamic design; Flow separation; UAV
Schlüsselwörter: Tragflügelprofil; Aerodynamischer Entwurf; Strömungsablösung; Drohne
* Correspondence and reprints.
E-mail address: dirk.schawe@dlr.de (D. Schawe).
1270-9638/02/$ – see front matter © 2002 Éditions scientifiques et médicales Elsevier SAS. All rights reserved.
PII: S1270-9638(01)01127-0

44 D. Schawe et al. / Aerospace Science and Technology 6 (2002) 43–51
Nomenclature
A Aspect ratio [–]
cd Airfoil section drag coefficient [–]
CD Drag coefficient [–]
CD0 Zero-lift drag coefficient [–]
cl Airfoil section lift coefficient [–]
CL Lift coefficient [–]
cm0 Pitching moment coefficient [–]
CP Pressure coefficient [–]
d Diameter of the propeller [m]
D Drag [N]
e Oswald’s efficiency factor [–]
g Acceleration of gravity [9.81 m s−2 ]
H Altitude [m]
L Lift [N]
m Takeoff weight [kg]
1. Introduction
Condor [4] rang in a new type of unmanned aircraft
flying fully autonomous from takeoff through landing in
high altitudes over long endurances. On 9 October 1988
Boeing’s Condor took off the runway for its first flight.
During seven more successful flights Condor achieved two
Ma Mach number [–]
n Revolutions per minute [1 min −1 ]
p Pressure [Pa]
P Power [W]
r/R Relative propeller radius [–]
Re Reynolds number [–]
S Wing area [m 2 ]
T Thrust [N]
v Velocity [m s −1 ]
x/c Relative chord length [–]
α Angle of attack [ ◦ ]
β Geometric blade pitch angle [ ◦ ]
η Propeller efficiency [–]
π 3.1416 [–]
ρ Density [kg m −3 ]
world records, one for an altitude of 20 416 m, and one
for an endurance of 58 hours 11 minutes. Condor was an
all-bonded composite aircraft with a 60.96 meter wingspan
of aspect ratio 36.6 and a lift-to-drag ration (L/D) of 40
operating continuously at lift coefficients up to 1.35 with
laminar flow over 50% of the upper and lower wing surfaces
at Reynolds numbers of 1 million. Condor is powered by two
Fig. 1. History of manned and unmanned high altitude and long endurance aircrafts.

130.5 kW turbocharged and liquid cooled six cylinder piston
engines.
Condor’s as well as other aircraft designs show that aerodynamics
plays an important role for operating manned or
unmanned aircrafts in high altitudes for extended periods of
time (Fig. 1). The aerodynamic design of such a vehicle is
greatly complicated because of the dramatic atmospheric implication
at an altitude of e.g. 24 000 m (see Table 1). Ambient
density and pressure are just a vanishing fraction of
their sea level values. The lower speed of sound increases local
Mach numbers, and higher kinematic viscosity decreases
Reynolds number. This denotes, that airfoils for HALE-
UAVs operate on fairly high lift coefficients (CL ≈ 1.2−1.5)
at relatively low Reynolds numbers (≈ 1.0 × 10 6 )andrelatively
high Mach numbers (≈ 0.4 − 0.6). For these conditions
we are not able to draw upon already in the past developed
airfoils. Sailplane airfoil data are very near, but are
designed for operating on low Mach numbers (≈ 0.05−0.2).
They might serve as starting geometries for computational
airfoil design and optimization procedures ([1,2]).
Another important issue for operating aircraft in the
stratosphere, is the selection of the power plant. Turbofan
engines are reliable, of compact size and therefore easy to
integrate into the aircraft airframe, with high acquisition
costs, low maintenance efforts but moderate operation costs
and high climbing performance. Unfortunately the thrust decreases
proportional with the density, i.e. unmodified turbofans
are limited to about 20 000 m. On the other hand
piston engines with all their accessories (compressors, turbochargers,
heat exchangers, two-stage gear box, propeller,
etc.) have a low but constant performance with increasing
altitude but are much more difficult to integrate. The maximum
altitude for turbocharged and liquid cooled piston engines
is about 26 000 m. The propeller layout for high altitude
operation must provide the aircraft with enough thrust
till density ratios of 1: 28.
2. Strato 2C – introduction
Strato 2C is designed as an instrument for ozone and
climate research which was a programme of the German
Ministry of Education and Technology (BMBF). The project
coordinator was DLR (German Aerospace Center) who was
responsible for the project management, the flight operations
and the scientific instrumentation and mission preparation.
The construction was commissioned 1992 by the German
company Grob.
Its main capabilities have been devoted to:
• research of the dynamics and chemistry of the atmosphere;
• environmental research (e.g. pollution produced by air
traffic);
• exchange process between troposphere and stratosphere;
D. Schawe et al. / Aerospace Science and Technology 6 (2002) 43–51 45
Table 1
Parameters of the International Standard Atmosphere for an altitude of
24 384 m (80 000 ft)
Variable Value Comp. to S.L.
Altitude 24 384 km (80 000 ft) –
Density 0.04353 kg m−3 3.6%
Pressure 2716.62 Pa 2.7%
Temperature 221.03 K 76.8%
Speed of sound 298.04 m s−1 87.6%
Kin. viscosity 3.26 × 10−4 m2 s−1 2 227%
Fig. 2. Dimensions of the Strato 2C.
Fig. 3. Strato 2C on its maiden flight on 31 March 1995.
• observation of land, ocean and polar icecaps from high
altitude.
The aircraft was designed for altitudes up to 24 000 m
and long endurance and range operations (18 000 km) in the
stratosphere. It was supposed to carry a scientific payload of
800 to 1 000 kg depending on the flight mission. Strato 2C is
fully built of fiber composite materials with a takeoff weight
of about 12 000 kg and a wingspan of 56.5 m and a length
and height of 23.98 m and 7.76 m, respectively (see Fig. 2).
The first successful flight was on 31 March 1995 (see
Fig. 3). 29 test flights were scheduled successfully until
August 1995. At the last flight Strato 2C reached its
maximum ceiling at 18 500 m.

46 D. Schawe et al. / Aerospace Science and Technology 6 (2002) 43–51
Fig. 4. (a) Altitude design mission; and (b) long endurance mission.
3. Strato’s missions
The aircraft was designed with regard to two main
mission profiles. The high altitude mission intended for
stratospheric research has a range of 7 000 km with a cruise
endurance of 8 hours at an altitude of 24 000 m and 800 kg
mission payload (see Fig. 4(a)). The long endurance mission
has a total time of 48 hours over a range of 18 100 km in
18 000 m ceiling (see Fig. 4(b)).
4. Aerodynamics of Strato 2C
For Strato 2C an airfoil for operating in high altitudes was
designed without disregarding the requirements for start,
climb, descent and landing. The design Reynolds number
of 1.0 × 10 6 combined with a fairly high lift coefficient at
relatively high Mach number of Ma = 0.43 causes normally
laminar separation bubbles with turbulent reattachments on
smooth surfaces which produce additional pressure drag.
These unusual design criteria yield to a laminar airfoil –
named LH37 [7] – of 17.5% maximum thickness located
at x/c = 0.354 and a maximum camber of 0.040 located at
x/c = 0.385. The pitching moment is cm0 =−0.13, and for
lift coefficients between 0.3 and 1.5 low drag is achieved.
The relatively high airfoil thickness affects the structural
weight positively and enlarges the torsional stiffness of the
high aspect ratio wing.
The main goal of the airfoil design was to achieve high
lift coefficients while the drag remains rather low. High lift
occurs from high pressure differences between the upper
and lower surface of the airfoil, i.e. the upper surface
of the airfoil features high suction and the lower surface
overpressure over the whole chordlength. A low friction drag
will be reached by large laminar extends even at high lift
coefficients. The drawback of such a design goal is that
the airfoil will respond very sensitively on a poor finish
qualtiy and/or slightest contamination of the wing. This
effect is demonstrated in Fig. 5. The subsonic wind-tunnel
MUB located at the DLR in Braunschweig (Germany) was
equipped with a smooth and clean LH37-wing of constant
Fig. 5. Experimentally determined lift coefficient versus angle of attack for
Ma = 0.39 and Re = 1.1 × 10 6 .
chordlength in order to realize quasi-2D conditions. The
upper curve shows the lift coefficient distribution versus
the angle of attack for the unchanged wing. In a second
experiment the point of transition was set to 7% of the
chordlength on the upper and lower surface in order to
simulate a poor quality and/or contaminated wing. The result
was a significantly reduced performance of the airfoil caused
by fully turbulent flow and additionally by the beginning of a
trailing edge separation for angles of attack of α = 7.5 ◦ and
higher. The maximum difference gained by laminar flow in
units of the lift coefficient is 0.4 at an angles of attack of
α = 10 ◦ . The airfoil is non-critical because for increasing
angles of attack the lift does not breakdown.
The performance of the airfoil was calculated with the
two-dimensional viscous aerodynamic design and analysis
code ISES of Drela/Giles [1] and experimentally verified in
the subsonic wind-tunnel MUB. For a Mach number of 0.3,
cl = 1.5 and 7.5 ◦ angle of attack we received a very good
conformity in the cp-distribution of computer simulation and
experimental measurement. In this case transition through a
laminar separation bubble occurred on the upper surface of
the airfoil at x/c = 0.38 and on the lower surface at x/c =
0.79 (see Fig. 6). Keeping laminar flow on the airfoil’s upper
surface till 38% chordlength at such high cl’s is a reasonable
value. If the lift coefficient will be reduced to cl = 1.25 the
laminar extend could be held up till 45% chordlength.
Airfoils where the flow remains laminar over a wide
chordlength at high lift coefficients produce low drag over
a wide cl-range as shown in Fig. 7. The drag increase
of the airfoil LH37 received from wind-tunnel experiment
compared with a computer simulation for Ma = 0.39 and

Re = 1.2 × 10 6 is pretty small until cl = 1.5. For higher
lift coefficients the drag increases significantly but without
cl-breakdown, which indicates too that the stall behaviour
of the airfoil is noncritical. The shapes of both polars are
very similar although the simulated one achieved lower drag.
The lower drag values of the simulated one is caused by the
restricted mathematical description of the flow physics on
which the computer program is based.
Fig. 6. Theoretically and experimentally determined pressure distribution
for Ma = 0.3, cl = 1.5, Re = 1.0 × 10 6 and α = 7.5 ◦ .
Fig. 7. Drag polar for the airfoil LH37 for Ma = 0.39 and Re = 1.2 × 10 6 .
D. Schawe et al. / Aerospace Science and Technology 6 (2002) 43–51 47
Strato 2C has a slight dihedral (1 ◦ ) triple swept wing with
a straight trailing edge and is equipped with an aileron and a
conventional Schempp–Hirth flap.
For the empennage symmetric laminar flow Eppler airfoils
were selected.
4.1. Wing-engine nacelle interference
The engine-nacelle is bluntly mounted on top of the wing
without fillets/fairings. In such regions flow separations
occur because of an appearing cross-flow caused by the
pressure differences between the flow around the nacelle
and the suction side of the airfoil. Due to the pressure
gradients a cross-flow from nacelle to wing is caused. The
boundary layer of the wing in the near region of the nacelle
will be thickened and therefore susceptible to separation,
especially when the wing operates at high lift coefficients.
Furthermore a vortex developes in the intersection of the
wing and nacelle. The separation area is wedge shaped,
running on the upper surface of the wing from the pressure
minimum under an angle of 45 ◦ spanwise to the trailing edge
(see Fig. 8).
Several methods could possibly weaken or dampen the
affinity of separation. Most methods provoke increasing
the energy in the boundary layer against positive pressure
gradients:
• cowling the wing-nacelle intersection with fairings which
will reduce the pressure gradient after the maximum airfoil
thickness;
• modifying the power plant integration by positioning
the nacelle with the engine inlet on the underwing
(see Fig. 9(a)), where the pressure gradients are much
lower, and therefore the boundary layer responds less
sensitive on disturbances. An additional advantage is
the possibility of integrating the landing gear into the
nacelle (see Fig. 9(b)), and thereby avoiding the vortex
generating belly fairing;
Fig. 8. Flow separation area in the wing-nacelle intersection.

48 D. Schawe et al. / Aerospace Science and Technology 6 (2002) 43–51
(a)
(b)
Fig. 9. (a) Modified power plant integration; (b) integration of the landing
gear into the nacelle (proposed by IABG).
• feeding the decelerated fluid particles of the boundary
layer with additional energy by blowing air into flow
direction. The flow velocity will increase and thereby
the separation risk reduced;
• the boundary-layer suction could be performed in the
area of increasing pressure, where the retarded flow will
be sucked off, and the non-retarded flow layers will
build the new boundary layer which is more resistent
against flow separation;
• influencing the boundary layer with vortex generators
in order to exchange high-energy fluid particles of the
outside flow with the boundary layer.
Three of these five methods – fairings, vortex generators
and blowing out – for avoiding flow separation in the wingnacelle
intersection were investigated in the subsonic windtunnel
MUB on a model with scale 1: 22.4.
The surveyed vortex generators could not prevent the flow
from separation. A more successful and less expensive approach
was the modification of the wing-nacelle intersection
with different fairings. These experiments made obvious
that an important parameter for avoiding separations completely
is to lengthen the fairing behind the wing trailing
edge. The fairing shown in Fig. 10 eliminated the separation
completely. Another promising solution was blowing air into
the boundary layer at x/c = 0.4 with a velocity of 80 m s −1
(�p = 1.5 × 10 5 Pa). For this configuration the separation
Fig. 10. Modified wing-nacelle intersection with an optimized fairing.
vanished completely. For the real aircraft the exhaust gases
of the engine could be blown into the wing-nacelle intersection
with a volume flow rate of 30 ls −1 .
A more detailed investigation about the wing-nacelle interferences
including wind-tunnel experiments and methods
to reduce their drag contribution are given in [5].
5. Performance evaluation of Strato 2C’s propeller
The propulsion system (see Fig. 11) is a combination of a
300 kW piston engine with a gas generator which serves as
a turbocharger and provides an additional jet thrust of about
12% of the propeller thrust at design altitude (24 000 m).
Between each compressor stage the air is cooled by intercoolers.
The compressor has an overall compression ratio of
32 : 1 and provides charge air for the engine’s manifold inlet
at 24 000 m of sea level condition.
Each of the two engines mounted in large nacelles on
top of the wing drives a wooden Kevlar composite coated
variable-pitch five-blade, constant speed and six meter
diameter propeller, working in pusher mode. The propeller
is abnormally strongly twisted from β = 43 ◦ at the propeller
hub to −10 ◦ at the tip.
Fig. 11. Propulsion system of Strato 2C.

Fig. 12. Strongly twisted propeller of Strato 2C. The four airfoil sections
are taken into account for the evaluation.
Fig. 13. Reynolds number distribution of one propeller blade at four
altitudes.
The propeller was designed for operating in the stratosphere
from 12 000 m up to the maximum mission altitude
of 24 000 m. In other words, the propeller must cover a remarkable
Reynolds number range from 1.8 × 10 5 to far beyond
1.0 × 10 6 (see Fig. 13). For the performance evaluation
four discrete flight levels 12 000 m, 18 500 m, 22 000 m
and 24 000 m have been selected. Therefore one propeller
blade was discretized into four airfoil sections (see Fig. 12).
The pressure distributions and the drag polars for these four
airfoils were calculated with the two-dimensional viscous
aerodynamic design and analysis code ISES of Drela and
Giles [1] and published in [6]. The problem of calculating
the drag polars of various propeller slices are the very low
Reynolds numbers below 200 000 at relatively high Mach
numbers between 0.6 and 0.8. Up to now ISES still offers
the highest accuracy for such flow conditions. But its error
should not be underestimated, and 20% to 30% error should
be taken into account. The performance characteristics of the
propeller are calculated on the basis of the afore determined
aerodynamic characteristics with a computer program developed
by Hepperle [3]. This computer code is based on the
combined blade-element and momentum theory which accounts
for details of the propeller airfoil geometry and their
D. Schawe et al. / Aerospace Science and Technology 6 (2002) 43–51 49
aerodynamic forces in order to design and analyse an optimal
propeller for the selected application.
The main assessment criterion for propellers is whether
they are capable of providing the airplane with sufficient
thrust for level flight at these operation altitudes, and
furthermore if they have an adequate climbing performance
reserve in order to reach the maximum altitude in a finite
period of time. The basis for calculating the propeller
performance are the drag polars in various slices, especially
between 60% and 90% of the propeller radius where the
main thrust is produced. In Fig. 14 we observe that for
altitudes above operation altitude 2 (18 500 m, see also
Table 2) the drag for the 60% slice experienced a strong
increase, which indicates flow separation. In the 80% slice
the flow is still attached, but separates nearby operation
altitude III (22 000 m). These observations indicate that
the propeller is limited to altitudes between 18 500 m and
22 000 m. This presumption will be confirmed if we look at
the thrust of the propeller, the maximum power of the engine,
and the drag produced by the aircraft.
This minimal thrust necessary for level flight is approximately
equal to the drag D of the airplane flying at velocity
v:
D = ρ
2 v2SCD, with CD = CD0 + C2 L
, (1)
πAe
where S = 150 m2 .
CL is calculated using
CL = mg
ρ
2 v2S with m = 12 000 kg.
e is the Oswald’s efficiency factor which depends
only on the Mach number. A value of e = 1 indicates
an elliptical lift distribution. For the calculations e =
0.865 = const, A = 21.28.
However, in the incompressible theory for positive propeller
thrust TP it will be assumed that the flow through the
propeller disk is incompressible and irrotational. From this
theory we are able to calculate the perfect/ideal propeller efficiency
ηid as follows:
�
2P
v = ηid
πρd2 �1/3 , (2)
(1 − ηid)
with P the power output and d the diameter of the propeller.
Here, the thrust of a perfectly working propeller is
Tid = ηidP
. (3)
v
If we pursue the propeller thrust and the aircraft drag (remark:
propeller thrust given for one engine and aircraft drag
refers only to the half aircraft. The additional jet thrust of
about 12% of the gas generator is not considered.) for maintaining
level flight with increasing altitude in Fig. 15 and

50 D. Schawe et al. / Aerospace Science and Technology 6 (2002) 43–51
Fig. 14. Drag polar of the propeller airfoil section at 60% and 80% radius. The roman numbers (I, II, III, IV) indicate the four operation altitudes which are
specified in Table 2.
Table 2
Physical properties for four discrete operation altitudes. Comparison of the aircraft drag in level flight with the thrust provided
by the propeller. (Remark: Drag, Thrust and Power are referred to one engine without considering the 12% additional thrust of
the gas generator.)
Operation altitude I II III IV
Altitude H [m] 12 000 18 500 22 000 24 000
Density ρ [kg m−3 ] 0.30 0.11 0.064 0.047
Flight velocity v [m s−1 ] 62.2 101.3 131.3 153.8
Mach number Ma [–] 0.21 0.34 0.45 0.52
Lift coefficient CL [–] 1.35 1.37 1.42 1.41
Zero–lift drag coefficient CD [–] 0 0.023 0.025 0.026 0.027
Drag coefficient CD [–]
Twist angle β75 [
0.054 0.057 0.061 0.062
◦ ]
Advance ratio
26.0 43.5 48.5 64.0
v/nd[–] 1.088 1.772 2.066 2.418
RPM n [ 1/min] 572 572 636 636
Power output P [kW] 179 298 300 300
Efficiency η [–] 0.87 0.87 0.84 0.64
Perfect efficiency ηid 0.96 0.96 0.97 0.97
η/ηid 0.91 0.91 0.87 0.66
Comparison aircraft drag in level flight D with propeller thrust TP
D [N] 2 376.9 2 468.1 2 524.3 2 563.3
TP [N] 2 502 2 556 1 909 1 252
Fig. 15. Propeller thrust and aircraft drag versus altitude. (Remark: both
forces refer to one Propeller or in lieu of the drag, of the half aircraft.)
Table 2 we observe that for 18 500 m level flight with only a
very modest climbing capability could be ensured. However,
above 19 000 m the drag exceeds the propeller thrust. For
further climbing to higher altitudes additional power equal
the aircraft weight multiplied with the sine of the flight-path
angle (= climbing velocity/flight velocity) is necessary. The
engine however, already operates, at 18 500 m, at its power
limit of 300 kW. The relative efficiency of the propeller
η/ηid which indicates the performance rate of a perfectly
working propeller, drops rapidly from 91% in 18 500 m to
66% in 24 000 m.
6. Conclusion
Airborne platforms for altitudes above 20 000 m are until
nowadays a challenge for the aerodynamic design. Each
meter altitude for a given payload is hard-won through
optimizing each detail of the aircraft. This is aggravated
by the fact that these details are interdependent. The main
design challenges are:

• a laminar wing, producing drag as low as possible at
high lift coefficients;
• a drag minimized fuselage;
• a flow optimized wing-fuselage intersection with only
low pressure gradients in order to avoid flow separation;
• the selection of a suitable power plant, and its integration
into the airframe without interfering the flow too
much; and
• if the power plant is a piston engine the propeller should
be adjusted optimally to the shaft power of the engine
and the altitude range to be covered.
The research aircraft Strato 2C accomplished these requirements
for its design altitude of 24 000 m only with regard
to the wing design based on the airfoil LH37 and the
selection of a turbocharged piston engine as power plant.
The other points need improvement. The specific problems
in summary were:
• the belly fairing for the landing gear at the fuselage
generates high drag vortical flow;
• the wing was mounted in high wing configuration
without suitable fairings;
• the large nacelles are just put on the wing with sharp
intersections. No additional devices influencing the flow
in the intersection are envisioned. Large flow separation
areas were the result;
• the propeller together with the 300 kW piston engine
is not capable to bring Strato 2C in its current layout
into altitudes of 24 000 m. With reasonable climbing
D. Schawe et al. / Aerospace Science and Technology 6 (2002) 43–51 51
velocities 18 500 m was reached. Further climbing is
possible but not practical because of the vanishing
climbing rate.
The preliminary layout of such aircrafts is still afflicted
with errors which should not be underestimated. The reasons
are the lack of suitable and accurate tools on the theoretical
side (CFD codes) as well on the practical side (windtunnels)
to determine the aerodynamics of a body flying at
low Reynolds numbers with high lift coefficients and high
Mach numbers. Nevertheless the available tools point out the
trends with 20–30% accuracy.
References
[1] M. Drela, M.B. Giles, ISES: A two-dimensional viscous aerodynamic
design and analysis code, AIAA paper 87-0424, 1987.
[2] R. Eppler, D. Somers, A Computer Program for the Design and
Analysis of Low-Speed Airfoils, NASA-TM-80210, 1980.
[3] M. Hepperle, Ein Computerprogramm für den Entwurf und Analyse
von Propellern, Institut für Flugzeugbau der Universität Stuttgart, 1984.
[4] R. Johnstone, N. Arntz, CONDOR – High altitude long endurance
(HALE) automatically piloted vehicle (APV), AIAA paper 90-3279,
1990.
[5] Ph. Tölke, A. Quast, Untersuchungen zur Ablösung im Bereich
der Flügel-Gondel-Verschneidung am Forschungsflugzeug Strato 2C,
DLR-IB 129-96/36, 1996.
[6] G. Wichmann, H. Köster, Leistungsnachrechnung des Propellers des
Höhenforschungsflugzeugs Strato 2C, DLR-IB 129-96/31, 1996.
[7] G. Wichmann, C.H. Rohardt, P. Hirt, Kenndaten für Profile: Profil DLR-
LH37, Luftfahrttechnisches Handbuch – LTH, Band Aerodynamik AD
41102-24, 1998.