A computer program for determining the flight characteristics of light aircraft was developed. The parameters which were used in the computer program are defined. The accuracy of the system for various types of airfoils is analyzed and the airfoils for which the system does not provide adequate data are identified. The application of a computer program for predicting the fuselage characteristics is discussed. The assumptions and parameters of the fuselage characteristics program are explained. It is stated that the computer programs make it possible to determine the response of a light aircraft to a small disturbance given the geometric and inertial characteristics of the aircraft

Smetana, F. O.

English

NASA Technical Reports Server

North Carolina State University at Raleigh
Department of Mechanical and Aerospace Engineering
Raleigh, N. C. 27607
n m a 1 t  Final Report Covering the Period
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Frederick 0. Smetana, Professor el
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https://ntrs.nasa.gov/search.jsp?R=19740025313 2020-03-23T04:41:00+00:00Z
The work conducted under this grant may be viewed as a logical continua-
tion of the activities carried out under previous grants in this series. The
principal objective of the series is to package aspects of modern aircraft
technology as developed for the air transportation industry and the military
in a form suitable for use by the general aviation industry. A secondary
objective has been to rekindle faculty and student interest in aeronautics
after years of almost total concentration on space.
The series began with the preparation of a three-volume review (NASA
CR-1484, 1485, 1486, March 1970) of the NACA/NASA literature published from
i940 to 1968 for its pertinence to light aircraft problems. This effort
indicated to the authors that the area of stability and control offered the
most promise for the immediate introduction of modern estimation and analysis
techniques. Accordingly, they prepared a 409-page Contractor's Report
(CR-1975, March 1972) detailing the application of these techniques to
typical light aircraft designs. Computer programs for performing part of
the analysis tasks were also prepared and listings and, user Instructions
were included in the report. In a follow-on report, CR-2016, (May 1972)
they discussed methods for verifying these estimates through flight tests
and provided computer programs by which one could calculate stability de-
rivative values if given the aircraft geometry. These programs., along with
those given in CR-1975, make it possible to determine the response of an
airplane to a small disturbance given the geometric and inertial characteris-
tics of the airplane.
In CR-2272 (June 1973) computerized methods for treating point and path
performance of light aircraft were given. These methods permit one to fit
any arbitrary drag-speed curve and power-speed curve and obtain very accurate
performance data.
-2-
A review of the work to this point was given at the SAE Business Air-
craft Conference in April 1973 (Paper #730305).
Unfortunately, construction of a rigorous drag-speed curve is very dif-
ficult. In the past, for preliminary design purposes such curves were usually
constructed from quasi-analytical correlations of rather badly scattered flight
test data taken on similar aircraft. However, in the mid 1960's the major
airframe contractors and others began to develop large computer programs to
treat portions of this problem in an analytical fashion. Reports on these
activities began to appear in the literature inf 1970 and 1971. Some of this
work was done at NASA or was supported by NASA. As a result, it was reported
in considerable detail. Examination of the reports indicated that the pro-
grams often contained many options not needed for light aircraft work and could4)
therefore be shortened substantially for light aircraft work. While no single
program was available to do an adequate job on an entire airplane, sufficient
programs were available which, when modified, could be patched together to
enable one to treat an entire airplane. The modifications needed, however,
were not insignificant nor was the patching process simple. This situation
was responsible for the genesis of the work on the present grant.
Begun in August of 1972, the work first undertook a review, speciali-
zation, and simplification of the computer program described in NASA CR-1843.
As originally received from Langley, the program deck was dimensioned to
compute the lift, drag,and moment characteristics of a four-element, two-
dimensional airfoil by the method of distributed vortices on the airfoil
surface (for the inviscid solution), coupled with a momentum integral type
of boundary layer solution (to determine the viscous contribution to the
flow). The flap portions of the program, however, had been removed. The
modifications to this program consisted of
-3-
(1) reducing arrays to a size consistent with a single airfoil element.
(2) cleaning up other aspects of the programming to reduce storage
requirements and speed execution time by a factor 20.
(3) fixing the segments used to represent the airfoil at 65.
(4) changing the method used to represent the airfoil wake within the
inviscid flow formulation.
(5) doubling the Reynolds number used in the boundary layer calculations.
(6) removing one of the two turbulent boundary layer calculation routines.
(7) modifying the computed lift coefficient to account for the loss of
lift which occurs physically because of the effect of the wake on the net
circulation.
(8) modifying the computed drag coefficient to account for the fact that
the computations do not always give-a zero drag value in inviscid flow as
they should. (This is primarily a precision problem in the formation and
inversion of a 65 by 65 matrix; rather than sort out and correct problems
associated with taking differences between two large, almost alike, numbers,
the residual drag in the inviscid case was treated as a tare to be subtracted
from the results of subsequent calculations.)
These modifications resulted in improved predictions of airfoil charac-
teristics. As received from Langley, drag estimates at low angles of attack
were 60% to 100% high. The modified program gave results within 15% of those
obtained by experiment for the same cases. Agreement with experimental lift
values was within 6% and frequently much better for CL < 0.8 and airfoil thick-
nesses between 9% and 18% of chord. Modern low drag airfoils up to 21% thick
are also handled well but older airfoils such as the 2424 produce regions of
flow separation. The latter cannot be treated adequately within the framework
of the viscous flow theory used. As a result, the characteristics of airfoils
-4-
with concave surfaces cannot be estimated properly. Very thin airfoils: and air-
foils with flat surfaces also cannot be treated satisfactorily because they lead
to ill-conditioned matrices. This problem is inherent in the method used to
represent the flow about such bodies. The lift predicted by the Langley
version of the program was slightly greater than that given by the NCSU
version (a lift curve slope about 6% greater). Both versions, of course, over-
predicted the lift at higher CL s
The Langley version of the program contained a provision for correcting
the results to higher Mach number flight. This provision was retained in the
NCSU version. A limited check indicated that it gave useful results, using
the NCSU lift and drag coefficients, to about 80% of critical Mach number.
Ordinarily, complete wings are treated by representing them with a vortex
lattice. While quite rigorous, this approach has two principal disadvantages:
1. For equal accuracy, the inviscid representation is many times (7 20)
more complex than the representation of a two-dimensional airfoil.
2. None of the available vortex lattice programs are as yet combined
with three-dimensional boundary layer programs. Further, the latter are not
generally available for non-axisymmetric bodies.
Fortunately, NASA CR-1646 contained a computerized method called STALL
for converting experimental two-dimensional data to three dimensions following
the lifting-line representation of moderate-to-high aspect ratio, unswept
wings. This method was originally intended to enable one to predict the stall
characteristics of finite wings of this type; but because most light air-
craft have such wings, it also seemed ideally suited to converting 2-D theo-
retical results to 3-D in a far simpler fashion than the vortex lattice scheme.
Since the 2-D airfoil program now treats a single component only, the portion
of STALL dealing with flaps was removed. Also, since the boundary layer portion
-5-
of the airfoil program does not permit one to treat separated flows, that
portion of STALL dealing with aerodynamic characteristics near stall was
removed. The experimental input data in tabular form was replaced by a
curve fit of the output of the airfoil program. Comparison of the results
indicate that for input data of equal accuracy there is less than a 3%
variation between the results given by original version of STALL and the modi-
fied version. Thus any failure to given suitable three-dimensional lift
and drag predictions for low-to-moderate angles of attack are a result of
failures to compute the two-dimensional characteristics correctly, provided
of course that the wing planform conforms to requirements (moderate-to-high
aspect ratio and unswept). Four comparisons with experimental data for com-
plete wings gave excellent results for the entire procedure.
The foregoing was accomplished during the first year of the grant.
During the second year, attention was turned to predicting the fuselage
characteristics. The writer became aware of the availability of three
programs which could be used as a starting point in this work:
(1) A combined wing-body program at the Ames Research Center which
represented the fuselage by line-sources
(2) A combined wing-body program described in NASA CR-2228 which
represented the body by source distributions on the body surface
(3) A body only program from the Naval Ship Research and Development
Center which represented the body by a source distribution on the surface
and, in addition, calculated the flow streamlines on the body surface.
All three programs were acquired, mounted on the local computer, and
checked out for the test cases supplied. Further study indicated that the
program from the Navy would be easiest to use for the intended purpose besides
offering the advantage of giving the flow streamlines on the body surface.
-6-
None of the programs provided for drag computations (Except wing induced
drag). Also, no general three-dimensional boundary layer programs were found
to be available. Thus in order to get a fuselage drag prediction of any
kind the following approach was adopted:
(1) Since quasi-axisymmetrical bodies at zero angle of attack cause the
flow about them to exhibit very small components in the cross flow direction
and
(2) Since it can be shown that for such bodies the boundary layer flow
along streamlines is like two-dimensional flow except near the nose and tail
of the bodies,
(3) Assume that the angle of attack is restricted to the neighborhood
of zero and that
(4) The wall shear and displacement thickness along streamlines can be
computed by the same boundary layer routines as are used for the airfoil.
(5) The complete drag is obtained by integrating the skin friction
over the surface and by appending a wake body whose size is determined by
the boundary layer displacement thickness near the downstream end to the
physical body, recomputing the inviscid pressure distribution over the entire
physical body plus wake body combination, and integrating this new pressure
distribution over the physical body to obtain the form drag as well as the
body lift.
The procedure was found to require a great many quadrilaterals in
order to represent the surface of the fuselage adequately (~ 500 or more).
As a result, about 2 minutes are needed to compute the lift and drag of a
single isolated fuselage at a = 0. In order to insure that the input data
for such a lengthy calculation is correct, it is first submitted to the plot
-7-
routine described in NASA TMX-2374. It is then converted by another routine
to the required input form for the body program.
Computed results for streamlined bodies of revolution were within
about 20% of published experimental data and computed results for a
Cessna 182 fuselage appeared reasonable although no experimental data are
available against which to compare.
Listings of these modified programs along with user instructions are
given in a 500-page Contractor's Report which has been submitted to NASA
for publication. Also included in this report are:
(1) An extensive literature review discussing the chronology leading
to the development of the theoretical bases of the programs.
(2) A detailed exposition of these theoretical methods.
(3) A discussion of wing-fuselage interactions and wing-tail
interactions.
(4) A discussion of ways to iniegrate the foregoing to obtain lift,
drag, and moment predictions for complete configurations.
(5) Several simpler programs for computing the lift of symmetrical
airfoils from M = 0 to M = 1.0 and for a closed-form solution of the inviscid
pressures over prolate spherioids. These may be used to indicate approximately
the results to be obtained by the more complex programs.
(6) A supplementary bibliography listing nearly 200 very recent
references thought to contribute to an understanding of the problem of
predicting lift, drag, and moment of complete configurations by analytical
methods.
It is hoped that this Contractor's Report will be useful to university students
and aeronautical engineers generally as well as to those charged with the
-8-
task of implementing the computer programs for their companies. Release
of this report by NASA is expected in late spring 1975.
The wing program and the earlier stability and performance programs
have been supplied to the major firms in the general aviation industry
on a no-cost basis. These firms indicate that they have made frequent
use of the programs with quite satisfactory results. The university can
now provide the body program as well upon request.
The reader will recognize that during the current grant period it
has not been possible to determine adequately the validity of the complete
lift and drag predictions. The time required to complete the program modi-
fications was somewhat longer than expected. The cost of running the program
is sufficiently high that one does not run a case without considerable
forethough. An most importantly, there is a paucity of reliable experimental
data against which to compare the predictions. The latter condition,
hopefully, will be corrected during the follow-on work now beginning under
grant number NSG-1077. In this work, an advanced technology light twin
aircraft will be given extensive flight tests designed specifically to
provide reliable data for comparison with the predictions of the whole
series of computer programs developed to date. As part of this effort it
is expected that the analytical work done to date on wing-tail interference
will be programmed for computer solution so as to be included in preparing
the lift, drag, and moment estimates of the complete airframe. Also being
prepared is a computer program for a new method to extract lift, drag,
and power information separately from flight test data. The latter will
provide the comparisons with the lift and drag predictions. Performance
predictions will be made first with the estimated lift, drag, and power
data and then with the values extracted from the initial flight results.
-9-
CR-1975, as noted above, contains methods for predicting the values of
the various airplane stability derivatives. These methods are computerized
in CR-2016. Estimates made by these procedures will then be compared with
the stability derivative values extracted from flight test data by other
techniques described in CR-2016. Thus, the Contractor's Report to be issued
on the next phase of the work will provide extensive comparisons (for one
airplane only) between the predictive theory developed to data and experi-
mental values as well as new computer programs for flight test data reduction
and wing-tail interference prediction.
In addition to providing the general aviation industry with some more
modern analytical tools, the effort of the last six years has had its impact
as well in aeronautical engineering education. The original review work
involved 10 undergraduate students, three of whom stayed around for graduate
work to participate in succeeding phases of the work. Three other graduate
students have also become involved in the activity for varying periods of
time. CR-1975 has been used as a supplementary text in the senior flight
vehicle design course and CR-2272 has been used as a supplementary text in
the flight vehicle performance course, neither of which were taught by the
present writer. Faculty at the University of Illinois (Champaign-Urbana)
requested and were sent copies of the stability and performance programs
for use in their classes. It is to be expected that as the power of these
programs and their ease of use becomes more widely known, faculties in the
57 departments offering Aerospace Engineering degrees will begin to reexamine
their curricula with a view toward optimizing available instructional time
by concentrating on those aspects of fluid mechanics and dynamics under-
girding the computer design methods and by giving the student the opportunity
to run some of these or equivalent programs in the course of his own design
experience.
-10-
One cannot conclude this review without noting that the work over the
last six years has benefitted very greatly from the thoughtful comments
and suggestions of many people in the industry and at the Langley Research
Center, in particular the project technical officers: first, Mr. Joe Stickle,
and now, Mr. Harold Crane. Their contributions are hereby greatfully
acknowledged.


Development of computer programs to determine the aerodynamic characteristics of complete light aircraft

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