This document summarizes the design of flight test methods for a Cirrus SR20 aircraft to be used in an aeronautical engineering course. It describes revising an initial GPS airspeed calibration method to improve accuracy and reduce variability. Flight test results applying the revised method showed acceptable agreement between calculated and instrument airspeeds. A simplified level flight performance test using power-off glides is also proposed to determine drag characteristics with less complexity.

1. Flight design report only
aerospace engineering writing question and need an explanation and answer to help me
learn.
hi please see the hw attached to follow the instructions that given basically all that you need
to do that you open the zip file and run the codes see the data write about it a paper
following the AIAA paper that is attached as well please let me know if you have any
questions or concerns attached references if there is any.
Requirements: follow the roles | .doc file
47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida Flight Testing on
Cirrus Aircraft M. Mandziuk*, S. Yurk*, M. Grashik*, Tianshu Liu‡ Department of Mechanical
and Aeronautical Engineering College of Engineering and Applied Sciences Western
Michigan University, Kalamazoo, MI 49008 Abstract This paper describes a flight testing
project on a non-experimental Cirrus aircraft to design flight test methods for use in a
university aeronautical engineering curriculum. Methods have been tested for determining
level flight performance and airspeed calibration through standardized procedures.
Additionally, a simplified method for approximately determining level flight performance
using power-off glides is designed. Flight testing indicates acceptable results with less
complexity and reduced variability. 1. Introduction Flight testing involves the gathering of
data in order to accurately predict the performance characteristics of a particular aircraft
design [1]. Data gathered from flight tests may also be used to optimize the performance of
all airplanes of that same type. Flight testing is an important, final step in the aircraft design
process and the data can be used to validate computational performance predictions. The
Flight Test Engineering and Design course (AAE 459) at Western Michigan University
(WMU) provides students with practical insight to the design and analysis of in-flight
experiments [2]. Unfortunately, the experimental Cessna R182 normally used for the class
has become a significant financial burden due to the maintenance needed and the fact that
the College of Aviation of WMU (normally responsible for the maintenance and flight of the
R182) is adopting new Cirrus aircraft. The problem has arisen that the flight testing course
needs a way to collect data utilizing the new, non-experimental Cirrus aircraft. Since the
new aircraft platform is non-experimental, FAA regulations require that no modifications be
made directly to the aircraft, especially to the exterior [3]. This project attempted to solve
the problem by formulating new procedures that utilize a transportable data acquisition
system, possibly being applicable to any aircraft. This project involved the design of flight

2. test methods and procedures to evaluate the performance of the Cirrus SR20 aircraft. The
WMU MAE department needed to establish new test procedures for the AAE 459 Flight Test
Engineering course. The procedures integrated the Appareo flight data system with
Western’s new fleet of Cirrus SR20 aircraft. Constraints for the project were rather limited.
Economic, safety, and repeatability were the primary concerns. The flight test methods
were designed and evaluated while minimizing the amount of flight time and
‡ Associate Professor, Corresponding Author Department of Mechanical and Aeronautical
Engineering, G-220, Parkview Campus Western Michigan University, Kalamazoo, MI 49008
tianshu.liu@wmich.edu, 269-276-3426, Senior Member AIAA *: Undergraduate Student,
Department of Mechanical and Aeronautical Engineering, Parkview Campus Western
Michigan University, Kalamazoo, MI 49008, Student Member AIAA Copyright © 2009 by the
American Institute of Aeronautics and Astronautics, Inc. All rights reserved. AIAA Paper
2009-0571 1
47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida equipment required.
Test methods maintained a degree of safety both in operation and procedures (safety
concerning the pilot, passengers, environment, other aircraft, and the test platform). Lastly,
the final designs provide an easily repeatable process for accurate data collection and
analysis by students. Western Michigan University’s shift to the Cirrus SR20 platform
required the development of a new flight testing system for use in the AAE 459 course. The
team utilized new technologies and techniques in the design of flight test methods. The
Appareo system and Cirrus/Avidyne Glass Cockpit System provide the latest in compact
flight data recording and processing. This was also the first application of the Appareo
system to general aviation aircraft flight testing. The project possessed issues regarding
safety and legality of in-flight devices. Federal Aviation Regulations were examined for
conflicting design constraints. The safety of each flight and its corresponding procedures
were considered. The Flight Test Method Design project only assessed a particular aircraft
and data system. As such, its results are most useful only to researchers or students
utilizing similar configurations. Redesign would be required for other platforms or
applications. This project improved the capability of WMU to instruct undergraduates in
flight test engineering methods. Therefore, WMU graduates will be better equipped to
handle the practical design and analysis of modern aircraft. 2. Cirrus Aircraft as a Flight
Testing Platform 2.1. Cirrus vs. Cessna Previous flight testing at Western Michigan
University was conducted in a highly modified Cessna 182RG aircraft as shown in Fig. 1(a).
Under experimental certification, the aircraft was outfitted with a myriad of sensors and
recording equipment, including a free stream air sampler boom. This modification
capability, while useful to WMU flight testing initiatives, came at a high cost. (a)
(b) Figure 1. (a) Cessna 182RG aircraft and (b) Cirrus SR20 Western’s transition from the
Cessna 172 Skyhawk to Cirrus SR20 aircraft for flight training availed the use of a much less
expensive test aircraft. The aircraft would remain under normal certification however,
meaning that no additional equipment could be permanently mounted to the systems or
airframe. Despite this restriction, the Cirrus provides an excellent test bed for flight testing
with its digital avionics. Supplemented by a removable compact flight data recorder, the
aircraft is capable of performing most standard flight tests. AIAA Paper 2009-0571 2

3. 47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida Methods designed in
this project must comply with the requirement for no additional modifications to the Cirrus
Aircraft, as well as compliance with aircraft operation Federal Aviation Regulations (CFR
Parts 61, 91) [3]. While methods are designed to comply with these regulations, during the
tests the pilot is the final authority as to the conduct of the flight. Requirements for the
Flight Test Methods Design Project changed throughout the semester. The final
requirements were the delivery of flight test research results and method instructions for
the GPS Airspeed Calibration The Cirrus Aircraft is a single engine monoplane used by
Western Michigan University in the Aviation Science program (pilot training). Figure1(b) is
a picture of a Cirrus SR20 in flight and Figure 2 shows the three side views of the Cirrus
aircraft. The aircraft seats four and is made primarily of composite, creating a smooth finish
and preserving weight. This material use makes it a highly efficient and streamlined
general aviation aircraft (compared to other single engine monoplanes). The Cirrus is
powered by a 76” constant speed propeller, driven by a Continental Teledyne IO-360-ES
piston engine. This produces approximately 200 brake horsepower and yields a cruise
speed range of approximately 115 to 155 knots. The aircraft has a useable load of 912 lbs,
and with full fuel tanks, a range of approximately 785 miles. An uncommon feature for
general aviation aircraft, the Cirrus is equipped a Complete Airframe Recovery System
(CAPS) parachute. When deployed, the CAPS is designed to safely lower the aircraft to the
ground. Figure 2. Three side views of the Cirrus SR20 aircraft 2.2. Avidyne Flight Data
System The Cirrus aircraft excels as flight test platform in its avionics. As shown in Fig. 3,
the Avidyne primary flight display and multifunction display, in combination with the
communication radios and dual GPS units, provide digital readouts of aircraft performance
and operating information. Additionally, engine operation data such as cylinder head
temperatures and fuel flow rates can be downloaded from the on board recorder for later
analysis log maintenance or flight test personnel. For flight testing in particular, the digital
readouts of the Avidyne avionics provides increased accuracy over traditional analog dial
gauges, as well as AIAA Paper 2009-0571 3
47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida advanced
information normally not available, such as true airspeed and current winds at altitude
(derived automatically from the aircraft’s flight path). (a) (b) (c) Figure 3. (a) Avionics
in cirrus aircraft, (b) Avidyne multifunction display (MFD), and (c) Avidyne primary
function display (PFD) 2.3. Appareo Data Recorder Beside the Avidyne system, a
commercial flight recorder, the GAU 1000 (Geospatial Awareness Unit), developed by
Appareo Systems (www.appareo.com) can be used to additional data. Appareo's GS Flight
Recorder is based upon GPS technology by utilizing MEMS gyroscopes, accelerometers,
barometric pressure, and a solid-state compass. As shown in Fig. 4(a), the 2½" GAU is
compact, portable, and easy to take from plane to plane. Multipurpose mountings include
highly durable window suction cups or velcro for dash mounting. The unit is designed to be
portable and run for extended periods on a self-contained lithium ion battery. However, it
will also operate plugged in from an AC or DC power supply while simultaneously charging
the battery. It can offer the longest record times — up to 22 hours with 64 MB removable
memory card. The GS Flight Evaluator software provides multiple views of your precise

4. route, including a precise flight path, detailed U.S. satellite imagery, and flight AIAA Paper
2009-0571 4
47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida instruments. Figure
4(b) shows a synthetic vision of landing based on the data provided by the GAU 1000
recorder. Therefore, the position, ground speed, altitude, roll, pitch, yaw, orientation and
magnetic heading of an aircraft can be recorded during flight testing with the portable GAU
1000 recorder. Complemented with the flight data output from a Cirrus aircraft, a data set
may be sufficient for basic flight tests. However, since an airdata boom and elevator angular
sensors cannot be installed on a Cirrus aircraft, the true airspeed, angle of attack, yaw angle,
and elevator angle cannot be directly measured. Some alternative methods could be used to
estimate these missing data. ) (b) ) Synthetic vision of landing . Flight Test 1 – GPS
Airspeed Calibration test the pilot will fly three different headings; this project team used
magntrack (aFigure 4. (a) Flight recorder GAU 1000 installed in a cockpit, (b 33.1. Initial
Method To conduct thisetic headings of 090, 210, and 330, but the actual headings used in
the test should not have a significant effect on the test results, as long as they are sufficiently
different to negate the wind effects, thus the headings are left to the student to define. On
each heading, after the aircraft has stabilized, the ground track and groundspeed are
recorded utilizing the GPS unit and/or hand recorded from the flight instruments. Figure 5
shows typical flight tracks for airspeed calibration. Due to wind speed and direction the
track the aircraft follows differs from the heading flown. The ground speed should also be
different in each direction, except in the case of no wind. According to the reference 2, the
three headings are flown to form vectors, which, when aligned as shown in Fig. 6, absorb
the effect of the Wind Speed (WS) and the True Airspeed (TAS) can be solved. The TAS and
WS are assumed to be constant between headings for the calculations. This procedure
should be repeated for multiple power settings and altitudes. The test should result in
three vectors resolved from the GS (vector length) and ground (vector direction) at each
altitude and power setting. The vectors can be aligned with their heads all meeting at the
same point, as shown in the figure below, to solve for the TAS and Wind Speed (WS). The
tails of the vectors will form a three-point-circle. The radius of the circle is the TAS and the
distance from the center of the circle to the point at which the heads of the AIAA Paper
2009-0571 5
47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida vectors meet is the
WS. There are several ways in which TAS and WS may be determined, all of which assume
that the WS and TAS are constant for each power setting and altitude. The first is a
graphical method in which CAD software, or any drawing utility (even graph paper!), may
be used to arrange the vectors as shown and determine the necessary parameters. A
numerical method is used to determine the TAS from the three equations with three
unknowhere Vwx and Vwy are the components of the wind velocity and the x and y are the
coordinates on 3.2. Test Results g parameters were recorded. The test was performed twice
for reasons descr3.3. ethod Revision t for this test the TAS was also recorded from the flight
instruments as a referewns
22wy32wx322wy22wx222wy12wx1)TAS()Vy()Vx()TAS()Vy()Vx()TAS()Vy()Vx(=+++=+++
=+++ an arbitrary Cartesian system of the tails of the ground speed vectors. Since these

5. equations are nonlinear they need to be solved numerically, and thus computer code would
need to be developed to solve this system of equations. The followinibed in the “Method
Revisions” section. The results for TAS and WS were calculated using the home-developed
MATLAB function GPSas.m. The “Ave. TAS” column in the results of the second test refers to
the average true airspeed recorded from the flight instruments for each track. The percent
error is then the error of the calculated TAS compared to the average TAS. Figures 7-10
summarize the results at 3000 ft and 9000 ft in two flights. MOn the second flighnce. The
use of this reference assumes that the pitot static system (used for measuring airspeed) has
already been calibrated by the aircraft manufacturer. This provides a good reference for
comparison with the GPS measurements, and it also justifies the use of the onboard pitot
static system for measuring TAS in subsequent tests. The “Fuel Burned” column was
omitted from the second flight as well, since it was determined that this parameter was not
a factor that needed to be considered when performing this test. Also, note the variability in
the wind speeds between the two tests. The wind speed on the day of the first flight varied
between 50 and 60 knots at 3000 feet. The calculations hold the wind speed constant for
each track in order to calculate TAS at a particular power setting, thus it is recommended
that the test be conducted under low-wind conditions. This was the primary reason for
retesting this procedure. It was also deemed unnecessary to gather data at specific power
settings, but rather more important to record what power setting the aircraft was at and
achieve a good spread of power settings, thus the “Percent Power” column on the final data
recording sheet should be left blank, instead of numbered from 100% to 50% in increments
of 10%. AIAA Paper 2009-0571 6
47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida Figure 5. Flight
path “triangles” from GPS airspeed test. Faster GS is shown in red and slower GS in
yellow/orange Figure 6. Relationship between the GPS velocity vectors, wind velocity
vector and true airspeed Figure 7. Calibration test at 3000 feet MSL altitude in the first
flight AIAA Paper 2009-0571 7
47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida Figure 8.
Calibration test at 9000 feet MSL altitude in the first flight Figure 9. GPS airspeed
calibration results for 3000 feet MSL altitude in the second flight Figure 10 (continued)
AIAA Paper 2009-0571 8
47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida Figure 10. GPS
airspeed calibration results for 9000 feet MSL altitude in the second flight 4. Flight Test 2 –
Level Flight Performance (PIW-VIW Method) 4.1. Initial Method The aircraft’s level flight
performance is highly affected by the conditions it is being flown in. However, the drag
characteristics and Oswald efficiency are inherent to the airframe geometry, and do not
change with conditions. Therefore, when testing the aircraft to determine its performance,
it is extremely useful to develop a method that can be used no matter what the conditions
and still maintain consistent results. The power required for level flight is given by
VSKW2VSC21P23DORρρ+= This formulae depends not only the velocity, but also the air
density and weight. In order to absorb the air density effect, the equivalent airspeed
(σVVe=) is introduced, where σ is the density ratio SLh/ρρσ=, where ρh is the air density at
altitude and ρSL is the air density at sea level. Further, since the weight changes during

6. flight because of the fuel that is burned, the standard airspeed, VIW, is introduced
2/1)/(STeWWVVIW=, where WS is the standard weight of the aircraft at sea level. It is
equivalent to the maximum takeoff weight of the aircraft. This includes all passengers,
cargo, fuel, and empty weight of the aircraft. WT is called the test weight. This is the
standard weight of the aircraft with the weight of the fuel burned subtracted from it. The
amount of fuel burned is recorded every time a new datum point is recorded. Therefore, for
every data point recorded, there should be a new test weight. Accordingly, the generalized
power, PIW, is given by 2/3)/(sTRWWPPIWσ=. The power (PR) is proportional to the
brake horse power (BHP) by PR = ηpBHP, where ηp is the propeller efficiency. The brake
horse power is obtained from engine charts provided by the engine manufacturer and
engine data collected during the flight test. The data recorded during the flight test and the
values obtained from the engine charts can be used to calculate VIW and PIW. Figures 11
and 12 are the engine charts for the Teledyne Continental engine on Cirrus aircraft that
were used for this method [4]. The first chart uses the engine manifold pressure and engine
RPM to determine the engine horsepower. The second chart uses the fuel flow rate and the
fuel mixture quality to determine the engine horsepower. The Best Power mixture line was
used with this chart because the engine was set to run at best power. The values are given
in AIAA Paper 2009-0571 9
47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida horsepower for both
charts and, therefore, must be converted to foot-pounds in order to maintain consistent
units. Before conducting this test, the team must record the initial parameters of the
aircraft, including the aircraft’s empty weight, fuel weight, and the weight of all passengers
and cargo. Once the initial data has been recorded, the pilot is required climb to the
appropriate altitude, specified by the project team. When the altitude is obtained, the plane
should be trimmed for level flight. After the plane has stabilized, the initial conditions,
including pressure altitude and outside air temperature, should be recorded. The fuel
burned, indicated airspeed, engine manifold pressure, engine RPM, the power percentage of
the engine, the fuel flow rate, and the true airspeed must also be recorded. These values can
be recorded by hand since the aircraft’s avionics display all the required data. No automatic
data recoding system is required. After the data is recorded, the power to the engine is
reduced by decreasing the manifold pressure by 1 or 2 inches. The data is recorded once
again. The power should be reduced until the aircraft can no longer maintain level flight
without increasing velocity. This test should be conducted at multiple altitudes to obtain a
sufficient amount of data. The process to conduct this test is simple and easy to perform.
However, it requires a lot of flight time, which increases the cost of the test. Testing should
be conducted in smooth air and low wind if at all possible. Figure 11. Engine power as a
function of the manifold pressure for the Teledyne Continental engine [4] AIAA Paper 2009-
0571 10
47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida Figure 12. Relation
between the fuel flow and brake horsepower for the Teledyne Continental engine [4] 4.2.
Test Results The parameters found below were recorded. The engine data that was
recorded was used to obtain the engine’s horsepower from two different engine charts
provided by the engine manufacturer. These values were required to calculate the

7. generalized power as described by the test theory [1, 2]. The test was performed at three
different altitudes: 3000, 6000, and 9000 feet. However, the engine data obtained at 9000
feet could not be used due to a limitation of the engine’s performance charts. Therefore,
only the data obtained for 3000 and 6000 feet was analyzed. The data was used to calculate
standardized velocity (VIW) and generalized power (PIW). A power speed relationship was
used determined from PIW and VIW. With these generalized variables, the power-speed
relation is written as )VIW(SKW2)VIW(SC21PIWSL2S3SLDOρρ+= The clear advantage of
using the preceding equation is that the PIW-VIW relation does not explicitly depend on the
air density and testing weight. Thus, data of the power and speed obtained at different
altitudes and weights collapses into a single curve. By multiplying both sides of this
equation by VIW, a linear relationship can be created bVIWaVIWPIW+=4)())(( By
substituting the VIW and PIW values calculated earlier into this linear relationship, a plot
can be created and the slope (a) of the line and the y-intercept (b) can be determined. Thus,
the only AIAA Paper 2009-0571 11
47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida unknowns that
remain are CDO and K. These can be determined by using the following equations for a and
b and solving for CDo and K SC5.0aSLDOρ= and . Figure 13 is an example of a plot that was
created using the above process. The equation of the line is displayed on the plot showing
the values of a and b. The slope of the line and the y-intercept were found and used to
calculate the aircraft’s parasite drag and Oswald efficiency. The Oswald efficiency, which
depends on the aspect ratio (AR) of the wing and is related to the K, is calculated by
)S/(KW2bSL2Sρ=)K(AR/1eπ=. Figures 14-19 summarize the raw data and extracted major
results for 3000 and 6000 feet. y = 0.0046x +
3E+0605000000100000001500000020000000250000003000000001E+092E+093E+094
E+095E+09VIW^4(PIW)(VIW) Figure 13. Linear fit to determine the parasite drag
coefficient and Oswald efficiency 4.3. Method Revision As mentioned above, the data that
was recorded for 9000 feet could not be used due to the limitation of the engine charts
provided. Therefore, it is recommended that this test should not be conducted at altitudes
higher than 6000 feet. This test was conducted two different times. The first test was
completed by decreasing manifold pressure by 2 inHg. It was decided by the team that
sufficient data could not be obtained by decreasing power by this increment. Therefore, for
the second test, manifold pressure was decreased by 1 inHg. This provided enough data
points, improving the accuracy of the test. Due to limitations of the engine performance
charts, limited values were obtained, as shown by the grayed in boxes found in Figs. 15 and
18. Due to the inadequacy of the engine charts, the team decided to use another method
described in the next section. AIAA Paper 2009-0571 12
47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida Figure 14. Raw
data for 3000 feet MSL altitude Figure 15 (continued) AIAA Paper 2009-0571 13
47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida Figure 15. PIV and
VIW data for 3000 feet flight (a) (b) Figure 16. The determined parasite drag coefficient
and Oswald efficiency for 3000 feet MSL altitude, (a) Results obtained by using engine
manifold pressure and engine RPM chart, (b) Results obtained by using fuel flow rate and
best power chart Figure 17. Raw data for 6000 feet MSL altitude AIAA Paper 2009-0571

8. 14
47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida Figure 18. PIV and
VIW data for 6000 feet flight (a) (b) Figure 19. The determined parasite drag coefficient
and Oswald efficiency for 6000 feet MSL altitude, (a) Results obtained by using engine
manifold pressure and engine RPM chart, (b) Results obtained by using fuel flow rate and
best power chart 5. Flight Test 3 – Level Flight Performance (Glide Test Method) 5.1. Initial
Method The level flight performance parameters, CDo and e, may be obtained from
graphical and numerical analysis of a series of simple power off glides. As shown in the
force diagram in Fig. 20, the lift and drag of a powerless gliding airplane can be estimated.
Following the determination of the overall lift and drag from the glide angle and TAS, the lift
and drag coefficients may be calculated. Density (ρ) is taken as the average air density
through the glide, (Sref) the aircraft wing reference area, and VTAS the average true
airspeed. From this point, a graphical approach may be used to resolve CDo and e. The
method utilizes a plot of CD versus CL2 in order to solve the level flight performance
parameters (similar to the PIW-VW method). This is limited in that multiple points (glides)
are needed to fit an accurate line and obtain reasonable results. However, flight testing has
indicated that even a few points can result in reasonable accuracy. As shown below, the
intercept of the plotted best fitting line is CDo, the zero-lift drag coefficient. The Oswald
efficiency, e, can be estimated from the line slope. AIAA Paper 2009-0571 15
47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida Figure 20. Force
diagram in glide This test was developed as an alterative to the PIW-VIW Level Flight
Performance test. It is not dependant on engine charts obtained from the manufacturer,
and therefore it tends to be more accurate if enough data points are recorded. This test
involves flying to a desired altitude, cutting power, and descending until another specified
altitude. The project team decided to glide from 9000 to 4000 feet. Just like the PIW-VIW
method, the initial weights are recorded. The pilot then flies to the specified altitude. Once
the aircraft stabilizes and the initial conditions are recorded, the pilot reduces engine power
to idle. The pilot descends for about 1000 feet before any data is recorded so he can trim
the plane to glide at a constant rate at a predetermined indicated airspeed. The indicated
airspeeds used by the team were 90, 100, and 120 knots. After the initial 1000 feet descent,
the true airspeed, indicated airspeed, and fuel burned are recorded every few seconds
during the glide. Once the bottom altitude is reached, the final pressure altitude and the
total time for the glide are recorded. This process is repeated for the other two indicated
airspeed settings over the same altitude range. 5.2. Test Results Recorded data and
calculated results are shown in Figs. 21 and 22. Figure 23 shows a linear fit of CD as a
function of CL2 and the extracted parasite drag coefficient and Oswald efficiency. The
resulting performance numbers are very reasonable for single engine monoplanes
(research on empirical data indicated that such aircraft typically have drag values near
0.025 to 0.045 and Oswald efficiencies of 0.4 to 0.6). Due to aircraft availability, the test was
performed in a Cirrus SR22. The only differences between the SR20 and SR22 are a more
powerful engine and slightly larger wings. The drag and aerodynamic efficiencies are very
similar. AIAA Paper 2009-0571 16
47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida 5.3. Method Revision

9. The altitude range for this test must be adjusted depending on weather conditions. Initially,
this test was conducted between 3000 and 8000 feet. However, it was found that below
4000 feet on the day of the test, the air was too unstable and the aircraft was not able to
maintain a steady glide. It was decided to increase the altitude range to between 4000 and
9000 feet so the air was smoother. Also, only three indicated airspeed settings were used
for this test. This only provided three points for data linearization and analysis. To create a
more accurate plot, no less than five airspeed settings should be used. Due to weather
complications, the team could not conduct another test with at least five airspeed settings,
so the data from the first and only test were accepted. The method, even with only the three
glides, was successful in achieving reasonable results for the parasite drag coefficient and
Oswald efficiency. The results do not however account for propeller drag (wind milling in
the free stream). Future research at WMU should be conducted to eliminate this error.
Figure 21. Raw glide test data AIAA Paper 2009-0571 17
47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida Figure 22. Glide test
calculations
0.00000.01000.02000.03000.04000.05000.06000.07000.08000.00000.10000.20000.30000
.40000.5000CL squaredCD Figure 23. Linear fit of CD as a function of CL2 6. Challenges
Faced A number of interesting, unforeseen challenges came to light during the design and
evaluation of the flight tests. These challenges were faced and remedied by the design team
in a number ways. The first of these challenges was a general unfamiliarity with the Cirrus
aircraft and avionics. The platform is brand new, and has a new “glass cockpit” system, and
even those with AIAA Paper 2009-0571 18
47th AIAA Aerospace Sciences Meeting, 5-8, Jan 2009, Orlando, Florida piloting experience
within the group took time to read operating handbooks and review the location of
pertinent data during the preflight test period. The design group was completely new to the
area of flight testing, and an extensive study of literature was done in order to become
familiar with the basic theories and principles. Another major challenge was the weather.
Each test needed to be performed under VFR (Visual Flight Rules) conditions. This means
that it cannot be cloudy or raining on the test day, so careful planning was required when
scheduling the tests of the procedures. Another factor in the weather was discovered to be
the winds. High wind conditions make highly variable conditions, this was especially
apparent on the first test of the GPS Airspeed test. One more challenge was the availability
and readability of engine performance data. The data is available via engine charts, but the
accuracy of this data is based on the chart accuracy and scale, and information obtained in
this manner was found to be highly variable. Lastly, a general lack of ability to benchmark,
or cross-check, the obtained data was encountered. Information such as a particular
airframe’s parasite drag coefficient could only be verified by either the manufacturer
(proprietary information) or wind tunnel/CFD testing. Such testing was beyond the scope
of this project. Instead, results were compared versus each other and confirmed reasonable
based on available empirical data for general aviation aircraft. 7. Conclusions Three types
of tests have been designed, adapted and evaluated for the Cirrus SR20 aircraft, including
the test procedures specifically designed for students. All tests can be successfully
performed without the inclusion of expensive and complex data equipment. For each

10. method, recommendations are given for students performing these tests to reduce
variability and increase the ease of testing operation. The glide test method is found to be a
simple alternative to the PIW-VIW method for determining similar results with less
variability, but continued research could concentrate on how to eliminated propeller drag
and improve the test. This project has provided a useful baseline for continued teaching
and research in flight testing at Western Michigan University. References: [1] Kimberlin,
R. D., “Flight Testing of Fixed-Wing Aircraft,” AIAA, Reston, VA, 2003 [2] Liu, T. and Schulte,
M., “Flight Testing Education at Western Michigan University,” AIAA Paper 2007-0700, AIAA
Aerospace Sciences Meeting and Exhibit, Reno, NV, 2007 [3] FAA, “14CFR91: General
Operating and Flight Rules,” US Government Printing Office, Washington DC, 2006 [4]
Teledyne Continental Motors, “Engine Specification Report, Model IO-360-ES 210 BHP @
2800RPM,” 2006 AIAA Paper 2009-0571 19
45th AIAA Aerospace Sciences Meeting and Exhibit, 8-11, Jan 2007, Reno, Nevada Flight
Testing Education at Western Michigan University Tianshu Liu† and Micheal
Schulte‡Western Michigan University Kalamazoo, MI 49008 ⋅This paper gives a concise
description of the basic flight testing methods employed in the course “Flight Test
Engineering and Design” (AAE 459) and a summary of student projects conducted on the
experimental aircraft Cessna R182 at Western Michigan University. The topics include the
experimental aircraft and instrument, experimental design and planning, standard
atmosphere and parameters, level flight performance, airspeed calibration, climb
performance, takeoff and landing, stick-fixed neutral point, and phugoid. The development
of a portable flight-testing system complemented with the data output from the onboard
instrument in a Cirrus aircraft is briefly discussed. 1. Introduction Flight testing is the
process of gathering information (or data) which will accurately describe the capabilities of
a particular type of airplane, and which can be used to accurately predict and optimize the
use of all airplanes of that same type in future missions. Flight testing of research airplanes
constitutes the gathering of data in regions of the flight environment where little past
information has been obtained. This information is then used to design future airplanes
that can operate safely in this new environment. Flight testing is at the end of the aircraft
design process and is a unique part of it. Flight testing involves various engineering
disciplines such as aeronautical, mechanical, electrical and structural aspects all together,
forming a technical discipline itself. Although playing an important role in the design of
aircraft, flight testing is not included in traditional curriculum of aeronautical engineering
education, and few universities offer such a course. It is partially due to the high cost for
maintaining and operating a fully instrumented experimental aircraft for flight testing. The
senior course “Flight Test Engineering and Design” (AAE 459) at Western Michigan
University (WMU) was originally developed by Professor Arthur Hoadley during the
collaborations with NASA Dryden Flight Center. A Cessna R182 has been instrumented as
an experimental aircraft for flight testing (see Fig. 1). The textbook used for this course is
“Flight Testing of Fixed Wing Aircraft” by R. D. Kimberlin [1], while the book “Introduction
to Flight Test Engineering” by D. T. Ward [2] is used as an additional reference. This course
provides aeronautical engineering students with a practical insight to the design and
analysis of in-flight experiments. This experience includes the use of microprocessor-based

11. data acquisition systems and electronic sensor interface. Lab projects emphasize the pre-
test, flight, and post-flight phases of flight-testing with an emphasis on safety of flight issues.
Typical topics include the lift and normal force coefficients as a function of the angle of
attack, airspeed calibration using GPS, estimation of the parasite drag coefficient and
aircraft efficiency factor, control effectiveness; takeoff and landing performance, neutral
point and maneuvering point, and dynamic longitudinal stability characteristics. The course
has enjoyed a success in attracting students as a unique part of aeronautical engineering
education at WMU. However, as the time elapses, the accumulated problems such as
overhauling the aging engine of the Cessna R182 in addition to regular maintenance and
insurance impose a significant financial burden that cannot be covered by the laboratory
fee. The situation would further deteriorate as the College of Aviation of WMU that is
responsible to maintain and fly the Cessna R182 for the course adopts a new flight platform,
Cirrus aircraft, rather than the Cessna aircraft. Therefore, in order to overcome these
problems, the course has to be reformulated to make flight data acquisition unit more
flexible to use in a non-experimental aircraft and more independent from a particular
aircraft (e.g. the Cessna R182). The rapid progress in micro-electronic-mechanical system
(MEMS) makes a small portable flight data unit possible, including MEMS gyroscope,
accelerators, barometric pressure sensor and solid-state compass. Thus, the position,
ground speed, altitude, roll, pitch, yaw, orientation and magnetic heading of an
†: Associate Professor, Department of Mechanical and Aeronautical Engineering, G-220,
Parkview Campus Western Michigan University, Kalamazoo, MI 49008
tianshu.liu@wmich.edu, 269-276-3426, Member AIAA ‡: Graduate Assistant, Department of
Mechanical and Aeronautical Engineering, Parkview Campus Western Michigan University,
Kalamazoo, MI 49008, Student Member AIAA Copyright © 2007 by the American Institute
of Aeronautics and Astronautics, Inc. All rights reserved. American Institute of Aeronautics
and Astronautics Paper 2007-0700 1
45th AIAA Aerospace Sciences Meeting and Exhibit, 8-11, Jan 2007, Reno, Nevada aircraft
can be recorded during flight testing with the portable compact unit. Complemented with
the flight data output from the Cirrus aircraft (a regular, non-experimental aircraft) such as
the engine power, manifold pressure, density altitude and air speed, a sufficient data set can
be obtained for basic flight testing projects. As a result of collaboration with the College of
Aviation, this unique flight testing course can remain at a much lower cost in the curriculum
of aeronautical engineering education at WMU. Nevertheless, a disadvantage of using a
non-experimental aircraft for flight testing is that an airdata boom cannot be installed on
the aircraft such that the airspeed, angle of attack and yaw angle cannot be accurately
measured in the undisturbed incoming flow. Some approximate methods could be used as
an alternative to estimate these quantities. The objective of this paper is to provide a
systematical description of the course “Flight Test Engineering and Design” as well as the
associated equipment and the experiment designs. The lessons learned from the problems
accounted at WMU may be beneficial to other universities offering similar courses. First,
the paper describes the experimental aircraft Cessna R182 and summarizes the flight
testing projects conducted on this aircraft. It also provides a concise review of the relevant
flight testing methods for general aviation aircraft. Then, we briefly discuss a combined use

12. of a non-experimental Cirrus aircraft and a portable flight data system for flight testing. 2.
Experimental Aircraft Cessna R182 The experimental aircraft Cessna R182 has been used
for the flight testing course and flight testing research of the effect of icing on aircraft
performance. The registration number is N1817R and serial number is R18200571. The
engine maker/model is Lycoming O-540-J3C5D and serial number is L-20434-40A. The
propeller maker/Model is McCAULEY B2D34C214 and serial number is 785587. Table 1
lists the basic parameters of Cessna R182. Table 1. Basic Parameters of Cessna R182
Engine: LYC O-540-J3C5D 75% Cruise: 156 kts Wingspan: 35 ft Horsepower: 235 Stall: 50
kts Length: 28.33 ft Rec’md TBO: 200 hrs Range: 520 nm Height: 8.75 ft Std Fuel: 61 gal Ser
Ceiling: 14300 ft Empty Wt: 1782 lbs Max Fuel: 80 gal Takeoff: 820 ft Gross Wt: 3100 lbs
Landing: 600 ft AR: 7.5 Wing Area: 173.6 ft2 As shown in Fig. 2, the Cessna R182 is
equipped with an airdata boom installed at the left-wing tip, allowing measurements of both
pressure and the flow angles to be measured well in front of the wing where the
measurements are not influenced by the shape of the aircraft. The pitot-static system is the
basic measurement method for determining speed and altitude, including two pressure
measurements. Total pressure (or pitot pressure) represents the pressure being applied to
the front of the airplane as it moves through the air. It is measured by using a pressure
transducer to measure the pressure inside a forwarding-facing tube at the front of the
boom. Static pressure represents the undisturbed pressure of the atmosphere at the
altitude that the airplane is flying. It is measured by side-facing tubes or holes on the top
and bottom of the airdata boom. The static pressure measurement can be related directly
to the altitude that the airplane is flying. The difference between the pitot and static
pressure can be related through the Bernoulli’s equation to the speed of the airplane
through the air. Immediately behind the total and static pressure tubes on the boom are
two vanes that pivot freely on posts extending vertically and horizontally from the boom. A
transducer measures the position of these vanes relative to the boom centerline. The
resulting angles are angle of attack and angle of sideslip. Both are key measurements for
determining the stability of an airplane. As shown in Figs. 3 and 4, the on-board instrument
includes sensors for measuring temperature, dynamic pressure, atmospheric pressure and
manifold pressure, GPS, rate gyros, accelerometers, two video cameras installed on the
wings. There are also sensors for motion of the control surfaces. Miniature gyroscopes
(gyros) measure the rate of rotation about the three axes of an aircraft (pitch rate, roll rate,
and yaw rate). Accelerometers measure the linear acceleration along the same three axes
(fore and aft - X, sideways - Y, and up and down - Z). The three accelerometers and three
gyros are usually very carefully aligned and mounted near the aircraft's center of gravity,
often on the some mounting platform. Figure 5 shows the Euro-card rack, IOtech data
acquisition system, and Labtop for retrieving data. The data collected from the instrument
on the Cessna R182 are listed in a table shown in Fig. 6. A special load cell was designed by
Landan et al. [3] for measurements of the thrust, but the accuracy of the thrust cell has not
been validated. The instrument is powered by an on-board auxiliary power unit (see Fig. 7).
American Institute of Aeronautics and Astronautics Paper 2007-0700 2
45th AIAA Aerospace Sciences Meeting and Exhibit, 8-11, Jan 2007, Reno, Nevada 3.
Experiment Design and Planning The objectives for specific flight tests are first set,

13. including developing the applicable theory, establishing the equations to be used in data
analysis, determine the needed inputs needed to achieve objectives, and determining how
and when the inputs are to be determined. Data that should be known, but are not collected
in flight are the aircraft weight and balance for each configuration, aircraft dimensions,
winds, and other weather conditions. In the use of the data acquisition system in the Cessna
R182, a user configuration file is established and saved for the data to be collected. The
calibration constants for the parameters in the user configuration file such as angle of attack
and pressure transducers should be verified and compared with the master configuration
constants. The following steps are used for planning flight tests on the Cessna R182:
Establishing the pre-flight configurations: • Switch setting on power panel correctly; •
Equipment and instruments setup properly; • Computer booted; • Flight test software
running the correct user configuration file; • Selecting the operation mode and verifying
that sensors are responding; • Writing a checklist covering the above items. Establishing the
ground system configuration: • Turning on radio telemetry receiver; • Starting telemetry
software; • Selecting a unique file name for the flight; • Starting the ground program on one
of the network computers and verifying data reception from the aircraft; • Adding these
items to checklist. Establishing the in-flight procedures: • Developing a written in-flight
checklist: – Verify procedure for moving computer into aircraft and its configuration on the
engineer’s lap (cables, etc.); – Plan what data is to be collected in each data file and the name
for the file; – Is data to be collected in single dynamic burst or several steady state bursts? •
Developing data sheet for in-flight use: – Place for notes and remarks about each burst of
data collection; – Place to record actual conditions for each burst of data (airspeed, altitude,
etc.); – Place to record data inputs not taken electronically. Establishing the post-flight
procedures: • Closing any open data files; • Turning off telemetry transmitter power; •
Turning off telemetry receiver power; • Stopping telemetry software; • Removing any non
standard experimental equipment; • Adding these items to checklist. Establishing data
download procedures: • Converting all data files by burst and/or frame depending on need;
• Connecting flight test computer to network; • Transferring files to appropriate ground
computer; • Loading files into spreadsheet software for analysis. Performing a dry run
simulation: • Using the developed checklist to perform the dry run simulation; • Verifying
the checklist order and content through simulation; • Correcting checklist and procedure as
necessary; • Performing all the checklist tasks by each team member. Pilot briefing • Prior
to flight the pilot is to be briefed on each configuration necessary to gather the desired data;
• The pilot will verify which data points can be hit for the given weather and time available.
• A checklist for this briefing should be developed so no items are over looked. American
Institute of Aeronautics and Astronautics Paper 2007-0700 3
45th AIAA Aerospace Sciences Meeting and Exhibit, 8-11, Jan 2007, Reno, Nevada 4.
Standard Atmosphere and Atmospheric Parameters The pressure and density distributions
in the standard atmosphere for the altitude less than 11 km are modeled by the following
equations for the pressure ratio and density ratio 256.5SLp/)h(pθδ== and , (1)
256.4SL/)h(θρρσ==where the temperature distribution is given by
)1000/h(0226.01T/)h(TSL−==θ. (2) Here, the altitude h is in meters, and the subscript “SL’
denotes the conditions at the sea level. For , the atmospheric temperature is kept at 216.66

14. K, the pressure and density are given by km11h>(⎥⎦⎤⎢⎣⎡−−=cc1chhTRgexpδδ) and c/θδσ=,
(3) where 225.0c=δ, 752.0c=θ, , and is the universal gas constant with a convenient unit for
the standard atmosphere. Figure 8 shows the pressure, density and temperature
distributions normalized by the conditions at the sea level in the standard atmosphere.
Table 2 shows the standard atmospheric properties.
K66.216Tc=K/s/m97.286R221=Altitude is the vertical position of an aircraft with respect
to the earth’s surface. However, because the atmospheric properties vary with altitude, it
can also be considered as the vertical location of a specific value of temperature, pressure or
density. Thus, the altitude can be defined in different ways. The pressure altitude is the
height in the atmosphere at which a given value of standard pressure exists, while the
density altitude is the height in the atmosphere at which a given value of standard density
exists. The true altitude (or tapeline altitude) is the actual height measured by a tapeline.
The indicated altitude is the height read on the altimeter. Note that the pressure altitude is
the indicated altitude when an altimeter is set to 29.92 in Hg (the standard sea level
pressure). Table 2. Standard Atmospheric Properties From Eqs. (1) and (2), we have the
pressure altitude and density altitude in meters, respectively, )1(1025.44h19.03pδ−×=,
)1(1025.44h22.03σρ−×=. (4) All of these various altitudes are measured in feet from a
common datum plane. The true altitude is of interest to the pilot for purposes of terrain
clearance. Density or pressure altitude is of much more significance for performance
determination. The true altitude which corresponds to a given density altitude may vary
considerably from day to American Institute of Aeronautics and Astronautics Paper 2007-
0700 4
45th AIAA Aerospace Sciences Meeting and Exhibit, 8-11, Jan 2007, Reno, Nevada day.
However, aircraft performance is always the same at the same density altitude. For
example, referring to Fig. 9, we assume the aircraft is flying at a true altitude of 5000 feet
above sea level, where the pressure is 23.98 in Hg and the temperature is 53 F. The
pressure is 23.98 in Hg at 6000 ft in the standard atmosphere. Thus, the pressure altitude is
6000 ft. A temperature of 53 F, together with a pressure of 23.38 in Hg, yields a density of
0.001928 slugs/ft3. This is the density found at 7000 ft. Thus, in the standard atmosphere
the density altitude is 7000 ft. The performance of the aircraft at 5000 ft on this day is
comparable to that of the same aircraft at 7000 ft on a standard day. Another altitude,
which is of considerable importance to the pilot, is the absolute altitude that is the height
above the terrain immediately below or above ground level. In the above example, Figure 9
indicates that the absolute altitude is 3000 ft. 5. Level Flight Performance The power
required for level flight is given by VSKW2VSC21P230Drρρ+=, (5) depending on not only
the velocity , but also the air density and weight. The factor K is related to the Oswald
efficiency by . In order to absorb the air density effect, the equivalent airspeed
1)eAR(K−=πσVVe= is introduced, where the density ratio SL/)h(ρρσ= is given by Eq. (1).
Further, since the weight changes during flight, the standard airspeed VIW is introduced,
i.e., 2/1STe)W/W(VVIW=, (6) where and are the weight in the testing condition and the
standard weight, respectively. The standard weight is typically the maximum take-off (T-O)
weight at the sea level. Accordingly, the generalized power PIW is
TWSW2/3STr)W/W(PPIWσ=. (7) With these generalized variables, the power-speed

15. relation is written as )VIW(SKW2)VIW(SC21PIWSL2S3SL0Dρρ+=. (8) The clear advantage
of the use of Eq. (8) is that the PIW-VIW relation does not explicitly depend on the air
density and testing weight. Thus, data of the power and speed obtained at different
altitudes and weights collapses into a single curve. According to Eq. (8), there is a linear
relation . An application of this relation is the determination of the parasite drag coefficient
and the Oswald efficiency factor based on measurements of the power required and the
corresponding airspeed in level flight. The power is proportional to the brake horse power
BHPb)VIW(a)VIW)(PIW(4+=rPr, i.e., rprBHPPη=, where pη is the propeller efficiency.
Similarly, with the generalized variables, the lift-drag ratio can be written as
2SLS2SL0DST)VIW(SKW2)VIW(SCW21WDρρ+=. (9) For level flight in which the thrust is
balanced by the drag, i.e., TD=, Eq. (9) is re-written as a convenient form for linear
regression to determine the parasite drag coefficient and the Oswald efficiency factor
SWKW2)VIW(SCW2W)VIW(TSLTS4SL0DST2ρρ+=, (10) when the thrust and airspeed are
measured. Sample Projects A typical project on the Cessna R182 is to determine the
parasite drag coefficient and Oswald efficiency factor [4]. The quantities acquired during
flight were the dynamical pressure, thrust from the load cell, angle of attack, pitch angle,
and ground speed. The actual data sampling began once the aircraft had reached level flight
at 8000 ft. The pilot adjusted the throttle to achieve calibrated airspeed of 130 kts, and then
decreased the speed by 10 kts until the lowest speed 50 kts was reached. The test runs
included the cases where the flap and landing gear were fully deployed. The total time for
the experiment was about 60 minutes. Figure 10(a) shows as a function )VIW(TAmerican
Institute of Aeronautics and Astronautics Paper 2007-0700 5
45th AIAA Aerospace Sciences Meeting and Exhibit, 8-11, Jan 2007, Reno, Nevada of and a
linear fit. Note that the data presented here are reprocessed from the original data.
According to Eq. (10), this result gives the parasite drag coefficient 4)VIW(123.0C0D= and
the Oswald efficiency factor . The parasite drag coefficient given here is much larger than
that for general aircraft while the Oswald efficiency factor is significantly lower. The causes
for these differences have not been clarified. This may be due to inaccurate measurements
of the thrust by using an unproven load cell [3]. For comparison, a plot of (PIW)(VIW) vs.
for a Cessna U3A is shown in Fig. 10(b) [5]. This gives the parasite drag coefficient and the
Oswald efficiency factor . Clearly, the determination of the parasite drag coefficient and
Oswald efficiency factor is sensitive to the accuracy of data.
21.0e=4)VIW(028.0C0D=175.0e=The lift coefficient and normalized lift coefficient of the
Cessna R182 in level flight were measured [6]. In level flight, the lift coefficient is and the
maximum lift coefficient is . The normalized lift coefficient is defined as , which is directly
related to the stall margin characterized by . The velocity was measured at each angle of
attack (AOA) for a given deflection angle of flap, and therefore the lift coefficient was
estimated. Figure 11 shows the lift coefficient and normalized lift coefficient of the Cessna
R182 as a function of AOA for several deflection angles of flap.
SV5.0/WC2Lρ=SV5.0/WC2stallmaxLρ=22stallLNLLNV/VC/CC==)C1(100smLN−= 6.
Airspeed Calibration A pitot-static system is used to measure both the airspeed and altitude,
as shown in Fig. 12. An airspeed indicator (ASI) measures the differential pressure while an
altimeter measures the absolute pressure although both use the same static pressure.

16. According to the Bernoulli equation, the equivalent airspeed is
⎥⎥⎦⎤⎢⎢⎣⎡−⎟⎟⎠⎞⎜⎜⎝⎛+−=−11pq)1(p2V)1/(cSLγγργγσ, (11) where is the difference
between the total and static pressures and ppqTc−=γ is the specific heat ratio (1.4 for air).
For incompressible flow, SLc/q2Vρσ= if mph200V< and ft15000hp