A&AE 251 Introduction to Aerospace Design
Fall 2002 Semester Design Project
Executive Summary
In the fall of 2002, a request for proposal was issued for an aircraft capable of flight on Mars. The proposal requested information on the launch vehicle and spacecraft necessary to transport the aircraft to Mars. The purpose of the aircraft was to survey and collect data from a large area of the Martian surface, specifically the Valles Marineris. It is believed evidence of water may exist in this canyon.
The final design has the ability to meet the established requirements while maintaining a reliable, relatively low mass structure with a high probability of success. This design is the Aries Alpha. The aircraft is an adaptation of the Darkstar UAV—utilizing high aspect ratio wings for increased loiter time. In addition, these wings slow the sink rate of the aircraft. Also, high aspect ratio wings perform better than most other wing designs at low densities, a major benefit in the low density Martian atmosphere.
The Aries Alpha is also a tailless aircraft. Large control surfaces with a proven operational capability will help to account for the stability provided by a tail. The most significant advantage of a tailless aircraft is the weight savings. Since weight is a critical factor in the cost of the mission, a tailless configuration is favorable in this design. The combination of the high aspect ratio wings and tailless fuselage provides a comparatively low amount of drag. As a result, the aircraft will be more efficient and require less propellant during its flight. This in turn will again lower the weight (cost) of the aircraft.
Unlike the Darkstar, the fuselage of this aircraft is not circular. Instead, the fuselage is cylindrical in appearance, although it has an elliptical cross section. There is sufficient volume in the fuselage to house the avionics, scientific instruments, engine and propellant. A German-manufactured Marquardt R43 rocket thruster was chosen as the propulsion system for this aircraft. The R43 has the lowest mass of any of the engines considered that met the thrust requirement. After the engine has completed its burn and the aircraft descends almost to the surface, the instrument pack will be jettisoned from the Aries Alpha and will land on the canyon floor, protected by an airbag system similar to Pathfinder landing system, only on a smaller scale. The data from the survey flight will be transmitted to Earth via the “Planned Mars Network.”
Considerations for the launch vehicle and spacecraft focused on minimizing cost and weight. The launch vehicle’s mass to GTO capability, combined with the size of the payload fairing, led to the choice of the Atlas V 551 Launch Vehicle System. This is the least expensive launch vehicle capable of performing this mission. Size and mass were also the primary factors in designing the spacecraft. For the interplanetary engine, the Pratt and Whitney RL-10B-2 was chosen because of its reliable performance and high specific impulse. An entry capsule protects the Aries Alpha during the flight from Earth to Mars and entry into the Martian atmosphere.
The landing sequence will closely follow the successful Mars Pathfinder mission. A combination of parachutes, thrusters, and airbag systems will be used to safely land the payload. Upon landing, the entry vehicle will right itself, using a staggered deflation sequence of its airbags to insure optimal orientation for aircraft launch. After the spacecraft doors unfold, the aircraft will take off with the assistance of a solid rocket booster. Once the aircraft velocity is above stall speed, the booster will jettison and the aircraft will begin its mission.
Conceptual bases for the spacecraft and aircraft systems in this proposal are based upon proven, successful platforms. Drawing from the Pathfinder and the Darkstar, the Aries Alpha will meet the requirements of the proposal and provide a successful mission.
Figure #1
Figure #2
Figure #3
Figure #4
I. Introduction 6
II. Mars Aircraft 7
1. Concept Selection and Sizing 7
1.1 Design Database 7
1.2 Concept Development and Evaluation 7
1.3 Selection of Concept 10
1.4 Carpet Plots 11
1.5 Selection of Aircraft Design Point 12
2. Aircraft Description 13
2.1 General Description 13
2.2 General Arrangement and Layout 14
2.3 Aircraft Model 14
2.4 Design Parameters and Specifications 15
2.5 Inboard Profile 16
3. Aerodynamics 17
3.1 Airfoil 17
3.2 Drag Build-Up 20
3.3 Thrust Requirement 21
3.4 High Lift Devices 21
4. Stability and Control 21
4.1 Center of Gravity 21
4.2 Trim Device Considerations 22
4.3 Tailless Stability Considerations 23
5. Propulsion 24
5.1 Engine Size and Thrust 24
5.2 Propellant Consumption 24
6. Performance 25
6.1 Rocket Assisted Takeoff 25
6.2 Flight Profile 25
6.3 Descent and Landing 26
7. Weights 28
7.1 Estimation Methods 28
7.2 Weight Breakdown 29
III. Spacecraft 30
1. Launch Vehicle 30
1.1 Launch Vehicle Database 30
1.2 Launch Vehicle Selection 30
1.3 Selected Launch Vehicle 31
2. Spacecraft Description 32
2.1 General Description 32
2.2 General Arrangement and Layout 32
2.3 Spacecraft Model 33
2.4 Spacecraft Launch and Payload Mass 33
3. Astrodynamics 34
3.1 Earth to Mars Trajectory 34
3.2 Atmosphere Entry Approximation Orbit 34
3.3 Required Delta V 35
4. Propulsion 35
4.1 Spacecraft Engine Selection 35
4.2 The Upper Stage Engine 35
4.3 Propellant Consumption 36
5. Spacecraft Mars Entry Sequence 36
5.1 Landing 36
5.2 Preparation for Aircraft Takeoff 37
6. Mass 38
6.1 Estimation 38
III. System Cost 40
1. System Cost Discussion 40
2. Comparison 41
I. Introduction
The AAE 251 Fall 2002 design project mission is to conceive an aircraft / spacecraft system that will bear a small package of scientific instruments to a region on Mars near the massive Valles Marineris canyon. The purpose of the mission is to search for evidence of the existence of liquid water and conceivably even life on Mars.
The significant design requirements for this mission are as follows: via a team selected launch vehicle, the spacecraft destined for Mars (which houses the Mars aircraft and its array of science instruments) is positioned in Geosynchronous Transfer Orbit (GTO) about the Earth. The spacecraft then exits GTO, passes through a heliocentric transfer orbit, and enters a predetermined orbit around Mars at an altitude of 500 km. Entering the Martian atmosphere and landing, the spacecraft has completed its part of the mission (except a possible role in relaying communications), having provided all the necessary propulsion for its voyage from GTO to arrival on Mars upon completion of its de-orbit burn. The method of Martian atmospheric entry is like that of the Mars Pathfinder mission: an aeroshell heat shield and parachute deceleration. Following entry, the fixed-wing Mars airplane departs the spacecraft (either during entry vehicle descent or following the landing) and climbs to an altitude of 0.5 kilometers at the best rate of climb, creating its thrust without the ability to use atmospheric oxygen. It then cruises 40 kilometers to the Valles Marineris at the best range velocity, at which point, upon reaching the canyon, the aircraft descends to an altitude of 0.5 kilometers below Martian “sea-level.” Flying at the best endurance velocity at this altitude, the aircraft uses the on-board scientific instruments to collect data from this region for a minimum of 45 minutes. Finally, the aircraft lands at the bottom of the canyon, descending with a sink rate no faster than 10 meters per second—using the last of its energy stores to continue measuring atmospheric conditions and transmitting data back to Earth.
In the overall system design, mass is used as a gauge of cost, and (of course) the least possible cost is preferred. Also, trade-offs must exist for the selected aircraft and spacecraft constituent concepts—for example, the means of propulsion, wing design, etc.
II. Mars Aircraft
1. Concept Selection and Sizing
1.1 Design Database
Prior to beginning the design
process, the team compiled a design database consisting of supersonic aircraft,
high-altitude aircraft, and Mars aircraft concepts. The supersonic designs were of interest as
the speed of sound in the low density Martian atmosphere is a relatively slow
237.5 m/s at “sea-level.” As a result, a
proposed Mars aircraft could feasibly exceed the speed of sound. Also, high-altitude aircraft were included
due to their low atmospheric density operating conditions. Flying on Mars near “sea-level” has been
compared to flying on Earth at an altitude of 100,000 feet. Other proposed Mars aircraft concepts were
investigated to yield insight into design parameters of aircraft intended for
operation on Mars.
1.2 Concept
Development and Evaluation
Prior to the team’s first meeting, each member brainstormed and sketched an aircraft for possible design consideration. Each aircraft was unique in its own way. Some designs heavily reflected characteristics of existing aircraft while others were developed completely from the imagination. Every aircraft was presented and described to the group for concept evaluation.
Concept A is a conventional airplane with a high aspect ratio mid-wing, conventional tail, and a single rocket engine. It has many characteristics similar to that of the Lockheed U-2. Disadvantages of this aircraft include its difficulty to fit in the payload fairing due to its high aspect ratio wings and the contamination its rocket engine would create.
Concept B is an interesting design of a joined-wing airplane with a canard and a single rocket engine. The upper wing is tapered with no sweep whereas the lower wing is swept with no taper. They are joined by endplates that serve as vertical tail surfaces. Drawbacks of this joined-wing design include its difficulty and precision required to construct (which is directly related to cost), the added difficulty in folding its pair of connected wings to fit in the payload fairing (if necessary), and the length of the fuselage required to provide a sufficiently large moment arm for the canard to trim the aircraft.
Concept C is a flying wing design. In order for this aircraft to maneuver, three variable thrust, gimbaled engines provide the needed pitch, yaw, and roll control. This aircraft could avoid the potential control problem of inadequate control surface input in the thin Martian atmosphere by using direct thrust for stability and flight control. The main disadvantages of this aircraft are its size (ability to fit in the payload fairing), the added weight from the three gimbaled-thrust engines, the amount of atmospheric pollution it could produce from its engines, and the high reliance on computer-controlled rocket engine gimbals to sustain stable, controlled flight.
Concept D is a high aspect ratio, mid-wing aircraft with a canard configuration and a conventional vertical tail. It is powered by a single rocket engine. The weaknesses of this aircraft involve its possible folding wings to fit in the payload fairing, the length of the fuselage needed for an effective canard moment arm, and atmospheric contamination.
Concept E is very similar to Concept D in that it has a high aspect ratio mid-wing with a canard configuration. The main difference is its means of propulsion. It is a propelled by two propeller engines, one in front, one in back; one is a pusher, the other is a tractor. The biggest disadvantage of this aircraft is the weight of its engine. This concept may also need folding wings to fit in the payload fairing.
Concept F is designed after the Darkstar Unmanned Aerial Vehicle. It is a tailless, high aspect ratio mid-wing rocket powered aircraft. Drawbacks of this concept include the atmospheric contamination produced by its engine and its possible need for folding wings.
Concept G is a high aspect ratio high wing aircraft with a canard configuration. Its vertical tail surfaces comprise an H-tail which adds to the aircraft’s stability, but increases its structural weight. Disadvantages of concept G include the disrupted flow over the main wing in takeoff due to the canard, added weight from the H-tail structure, and atmospheric contamination of the rocket engine.
Figure # 5 Figure #6
To evaluate each concept, the team created a list of objectives in order to rank the concepts. These objectives include: low mass, small spacecraft size, long loiter time, proven concepts, landing descent rate, atmosphere contamination, positive lift from trim devices, public relations image, ease of construction, cruise velocity, and landing survivability. They were then ranked against each other to determine which objective the team thought was most important in designing the aircraft (Table #1). The team then took individual votes to give a weight value to each objective when evaluating the concepts (Table #2).
Table #1
|
|
A |
B |
C |
D |
E |
F |
G |
H |
I |
J |
K |
Score |
|
A Low mass |
---- |
1 |
1 |
1 |
1 |
1 |
1 |
1 |
1 |
1 |
1 |
10.0 |
|
B Small size (spacecraft) |
0 |
---- |
½ |
1 |
1 |
1 |
1 |
1 |
1 |
1 |
1 |
8.5 |
|
C Long loiter time |
0 |
1/2 |
---- |
1 |
1 |
1 |
1 |
1 |
1 |
1 |
1 |
8.5 |
|
D Proven concepts |
0 |
0 |
0 |
---- |
1 |
1 |
1/2 |
1 |
1/2 |
1 |
1/2 |
5.5 |
|
E Landing decent rate |
0 |
0 |
0 |
0 |
---- |
1 |
1/2 |
1 |
1 |
1/2 |
1/2 |
4.5 |
|
F atmosphere contamination |
0 |
0 |
0 |
0 |
0 |
---- |
0 |
0 |
0 |
0 |
0 |
0.0 |
|
G Positive lift from trim devices |
0 |
0 |
0 |
1/2 |
½ |
1 |
---- |
1 |
1 |
1/2 |
1/2 |
5.0 |
|
H PR/image |
0 |
0 |
0 |
0 |
0 |
1 |
0 |
--- |
0 |
0 |
0 |
1.0 |
|
I construction ease |
0 |
0 |
0 |
1/2 |
0 |
1 |
0 |
1 |
---- |
0 |
0 |
2.5 |
|
J Cruise velocity |
0 |
0 |
0 |
0 |
½ |
1 |
1/2 |
1 |
1 |
---- |
1/2 |
4.5 |
|
K Landing survivability |
0 |
0 |
0 |
1/2 |
½ |
1 |
1/2 |
1 |
1 |
1/2 |
---- |
5.0 |
Along with having an aircraft that could stay in the air a long time, an aircraft with a low mass and a small spacecraft size, which are both directly related to cost, were the three most important objectives in the team’s opinion to consider. Next to creating positive lift from trim devices, the team wanted an aircraft with a design that has been used and flight proven. The ability to survive a rough landing on the Martian surface was also desired in order for the scientific instruments to continue sending data after the mission was complete. As stated in the request for proposal, a low descent rate was necessary along with a high cruise velocity to quickly travel to the area of interest. At the end of the list came the ease of construction, the public relations image, and atmospheric contamination. The team believed that if a plane, even if it were difficult to build, was sent to fly on Mars, the public’s opinion of how the plane would look and the amount of pollution it would create were negligible compared to the engineering design considerations enabling the aircraft to make potentially breakthrough scientific discoveries.
Table #2
|
Relative
Weights |
||
|
Objective: |
Votes |
Decimal % |
|
Low mass |
83 |
0.276 |
|
Small size |
59 |
0.196 |
|
Long loiter time |
41 |
0.136 |
|
Proven concepts |
27 |
0.090 |
|
Positive lift from trim devices |
19 |
0.063 |
|
Landing survivability |
14 |
0.046 |
|
Landing decent rate |
10 |
0.033 |
|
Cruise velocity |
17 |
0.056 |
|
Construction ease |
16 |
0.053 |
|
PR/image |
13 |
0.043 |
|
Atmospheric contamination |
1 |
0.003 |
Table #3
|
Concept
Evaluation |
||||||||
|
Objective |
Weight |
A |
B |
C |
D |
E |
F |
G |
|
Low Mass |
0.276 |
0 |
-3 |
3 |
0 |
0 |
3 |
0 |
|
Small Size |
0.196 |
-3 |
3 |
3 |
-3 |
-3 |
0 |
0 |
|
Long-Loiter |
0.136 |
3 |
0 |
-3 |
3 |
3 |
6 |
-3 |
|
Proven Concepts |
0.090 |
3 |
-3 |
0 |
0 |
0 |
0 |
0 |
|
Positive Lift From Trim Devices |
0.063 |
-3 |
3 |
0 |
3 |
3 |
0 |
3 |
|
Cruise Velocity |
0.056 |
3 |
6 |
6 |
3 |
0 |
3 |
0 |
|
Construction Ease |
0.053 |
6 |
-3 |
-3 |
3 |
3 |
0 |
0 |
|
Landing Survivability |
0.046 |
0 |
0 |
0 |
0 |
0 |
0 |
0 |
|
PR/image |
0.043 |
0 |
6 |
6 |
3 |
0 |
3 |
3 |
|
Landing Descent Rate |
0.033 |
3 |
-3 |
3 |
0 |
0 |
3 |
0 |
|
Atmospheric Contamination |
0.003 |
-3 |
-3 |
-6 |
-3 |
6 |
-3 |
-3 |
|
Score |
|
0.477 |
0.006 |
1.524 |
0.327 |
0.186 |
2.031 |
-0.288 |
1.3 Selection of
Concept
Using these objectives and assigning a number rank (6 being “the best” and -6 being “the worst”), the team found that concept F, modeled after the Darkstar UAV, and concept C, the flying wing, best fit the design requirements. Upon further inspection and a final decision, concept F was chosen as the teams’ aircraft for design.
Table #3
Concept F is based on the Darkstar UAV because it had mission objectives similar to those described in the request for proposal. These include low atmospheric density flight conditions, long endurance flight, and autonomous robotic flight. With its blended wing body, high aspect ratio wings, and tailless configuration, the team believed that it would be a good design to use. A single rocket engine is thought to easily provide enough thrust for the minimum cruise velocity while a shortened fuselage (which leads to less empty weight) and a smaller spacecraft payload strengthen its case against the runner-up, concept C.
1.4 Carpet Plots
The aircraft sizing process centered on the creation and interpretation of carpet plots to choose the values for the design parameters of wing loading, thrust-to-weight ratio, and aspect ratio. Plots of takeoff weight versus wing loading were made for aspect ratios of 14, 15, 16, and 17, each with four thrust-to-weight ratio curves (0.1547, 0.221, 0.2873, and 0.3315). On top of these plots, constraint curves were added in order to place the appropriate limits on the aircraft design space to ensure the aircraft’s capability to meet mission requirements. The two mission requirements which showed up as constraint curves on the carpet plots were the best range velocity of at least 160 m/s for the cruise-climb and the sink rate of less than 10 m/s for the final descent. These two curves did not intersect, nor did they tightly constraint the aircraft design point.
At the time of carpet plotting, the team had already selected the Eppler E 214 low Reynolds number airfoil for use on the Aries Alpha. After consulting with Professor Crossley, an estimated drag divergence Mach number of 0.72 was accepted for this 11% thick airfoil. The Aries Alpha, in order to fly efficiently at its best range velocity, must remain under the drag divergence Mach number in order to avoid the substantial increase in drag that occurs at that flight velocity. The team took this information, along with an estimated peak altitude for the aircraft’s cruise-climb of 1.5 km (based upon the performance of a similar sample Mars aircraft concept) to determine that the velocity corresponding to Mach 0.72 at 1.5 km altitude on Mars is 169.26 m/s. Realizing that the drag divergence velocity for the E 214 airfoil with zero quarter-chord wing sweep was less than 10 m/s above the required best range velocity, the team determined that an additional constraint curve should be added to the carpet plots. Doing this ensured that the aircraft flies below the drag divergence Mach number of its wings at its best range velocity.
Figure #7
1.5 Selection of
Aircraft Design Point
The aircraft design point location is motivated by several considerations, the most important of which is minimizing takeoff weight. The takeoff weight must be kept to a minimum while staying within the design space dictated by the mission requirement and drag divergence Mach number constraint curves. If the aircraft is to meet the minimum best range velocity of 160 m/s, the design point must be located to the right of the black best range velocity constraint curve on the carpet plots. Similarly, in order to have a sink rate less than 10 m/s on the final gliding descent of the mission, the aircraft design point needed to be to the left of the light blue sink rate constraint line. Finally, the team-defined constraint line ensuring that the best range velocity is below the drag divergence Mach number necessitates that the design point be to the left of the green constraint line on the carpet plot.
Simultaneously considering all three constraints and striving to minimize takeoff weight resulted in the design point being selected on the T/W = 0.1547 line just to the left and above the intersection point with the (green) drag divergence Mach number constraint line. The team selected design points on each of the four aspect ratio plots (14 through 17), and calculated the takeoff weight based upon the aspect ratio and corresponding wing loading value for each one. The results of this aspect ratio comparison were graphically interpreted in the form of a plot showing takeoff weight versus aspect ratio. The minimum value for takeoff weight occurred for an aspect ratio of 16, resulting in the team’s choice of 16 as the aspect ratio to use on the Aries Alpha.
Figure #8
One final run of the team’s sizing code produced the final
aircraft parameters for the conceptual design, which are detailed in a
following section.
2. Aircraft Description
2.1 General
Description
The final Aries Alpha design
is very similar to the Lockheed Martin Tier III- Darkstar Unmanned Aerial Vehicle. Like the Darkstar, it is a tailless
aircraft with long, high aspect ratio, unswept wings and a short fuselage. In order to account for the inherent
stability lost by choosing to not use a tail, the size of the control surfaces
on the wing was increased. In addition,
a computer system similar to that used on the Northrop Grumman B-2A bomber was
added to operate the control surfaces and keep the aircraft stable. This system is comprised of four
computers. Several times per second,
each of these computers decides if any corrections need to be made to the
position of the control surfaces to remain stable. If three of the four computers agree on the
same correction, electric motors move the control surfaces to the correct
position. In doing so, the aircraft
remains stable, even without a tail. As
previously stated, the benefits of this tailless design include a shorter,
lighter (and thereby cheaper), easier to launch aircraft with a decreased drag
coefficient. High aspect ratio wings
minimize induced drag and are efficient during the loiter period, which is the
key aspect of the data collection mission.
Minimized induced drag produces a high lift-to-drag ratio, which then
lends itself to a lower sink rate.
Having a tailless aircraft reduces weight and leads to a shorter
fuselage, both of these factors yield an easier fit in the payload fairing.
2.2 General
Arrangement and Layout
The aircraft wings look very much like the wings on a sailplane. Due to their extreme length, the wing support structure will be somewhat heavy. Therefore, no fuel or electronics will be stowed inside the wing. Instead, all necessary fuel and instruments will be carried in the fuselage. Not only will this reduce the weight of the wing, but it will also move the center of gravity forward on our aircraft, thereby increasing its stability.
The airfoil chosen had to be designed for flight in extremely low Reynolds
number conditions. After analyzing several airfoils designed for low Reynolds number flight, the Eppler E 214 airfoil was chosen. This design was determined to be the most efficient airfoil available that would provide enough lift in Martian atmospheric flight conditions. It provided the longest loiter time and the highest lift-to-drag ratio during cruise-climb. The team desired to use zero quarter-chord wing sweep for the wing in order to allow for easier folding of the wings and be more conducive to stowing inside the entry vehicle. Drag divergence Mach number became a significant concern with this straight wing design.
Finally, instead of using an “air-breathing” turbofan, like the one used by the Darkstar, a liquid bi-propellant rocket thruster will propel the Aries Alpha. Since the Martian atmosphere is composed primarily of carbon dioxide, no flight proven air-intake engine would be able to operate on Mars. A liquid rocket was preferred instead of a solid rocket because of the ability to throttle the thrust on the liquid rocket as needed. The main consideration in selecting a liquid rocket was finding a rocket that meets the thrust force requirement with the least possible mass. The thrust force requirement was based on the approximate weight of this aircraft and the thrust-to-weight ratio determined by the sizing code. After examining several thrusters, a scaled down version of the German made Marquardt R43 was chosen.
2.3 Aircraft Model
Figure #9
Figure #10
2.4 Design Parameters
and Specifications
Table #4
|
Aircraft Specifications |
|
|
Gross Takeoff Weight |
317.2 N |
|
Wing Loading |
167.5 N/m2 |
|
Thrust-to-Weight Ratio |
0.1547 |
|
Aspect Ratio |
16 |
|
Wing Area |
1.897 m2 |
|
Wing Span |
5.510 m |
|
Mean Aerodynamic Chord |
0.352 m |
|
Root Chord |
0.430 m |
|
Tip Chord |
0.258 m |
|
Quarter-Chord Wing Sweep |
0.0 degrees |
|
Taper Ratio |
0.60 |
|
Thickness to Chord Ratio |
0.1111 |
|
Fuselage Major Diameter |
0.7 m |
|
Fuselage Minor Diameter |
0.3 m |
|
Fuselage Length |
1.60 m |
The dimensions of the aircraft can be seen in the table, model, and schematic. In order to calculate the fuselage interior volume required to stow avionics, scientific instruments, and propellant, as well as the wetted area of the fuselage, it was approximated as a cylinder with rounded ends having diameter of 0.4 meters. However to in order to more closely resemble the original design concept of a blended wing-body aircraft, it was decided to use an elliptic cross-section for the fuselage rather than a cylindrical cross-section.
2.5 Inboard Profile
Aries
Alpha’s interior profile is configured as follows. Avionics have been
placed in the front portion of the aircraft. Just behind the avionics section
is the compartment storing the scientific instruments including the airbag
landing system. A panel door is positioned below the instruments, which will be
blown off during the landing sequence to allow for their deployment. The
nitrogen tetroxide and hydrazine fuel for the engine is kept behind the science
instruments and immediately in front of the engine. The fuel was chosen to be
stored in the fuselage and not in the wings for two reasons. First, having the
fuel cells close to the engine minimizes the pipes and pumps necessary to get
the fuel to the engine, thus reducing weight. Second, the type of fuel used
requires temperature regulation, which is more easily accomplished if the fuel
is stored in a central location rather than spread throughout the wings. Using
insulation and temperature control systems, the fuel will be kept at
approximately 2 degrees Celsius, which is above their freezing points of 1.5
and -11 degrees Celsius for hydrazine and nitrogen tetroxide respectively.
Stability was also considered in the aircraft’s inboard profile, positioning
the aircraft’s elements such that the center of gravity would be ahead of the
neural point.
Figure #11
3. Aerodynamics
3.1 Airfoil
In selecting an airfoil for the Mars aircraft, the team considered four NACA airfoils and seven additional “low Reynolds number” airfoils. Before collection of airfoil data such as lift, moment, and drag coefficients could begin, an estimate for the Reynolds numbers the aircraft’s wings would experience in flight was needed. This was accomplished using the parameters from a sample Mars aircraft, approximated densities and viscosities at the pertinent altitudes in the Martian atmosphere, and the predicted velocities of the sample aircraft during cruise and loiter. The computations (completed in a MATLAB script and functions) resulted in a range of expected Reynolds numbers from 84,000 for loiter to 125,000 for cruise. Since available published lift curve and drag polar data for the NACA 2415, NACA 4415, NACA 23015, and NACA 65(216)-415 airfoils includes curves corresponding to Reynolds numbers as low as 2.6-3.1 x 106, the team determined simply assuming the values are approximately the same as at Reynolds numbers on the order of 100,000 would be imprudent. As a result, research was conducted in the NASG airfoil database to add several “low Reynolds number” airfoils for consideration of airfoils which have been tested at Reynolds numbers near those the Mars aircraft will likely experience. This search resulted in the inclusion of the following seven airfoils in the airfoil selection process: E214, E387, MB253515SM, S4233, E205, S4180, and SD7037. The test Reynolds numbers for these airfoils range from 97,000 to 127,000, producing significantly more reliable data for use on a Mars aircraft.
The design team identified the following seven primary objectives in selecting the airfoil: high loiter L/D, low moment coefficient, high cruise L/D, gentle stall behavior, low angle of attack during loiter, high maximum lift coefficient, and low angle of attack during cruise. High loiter lift-to-drag ratio is crucial for efficiency in the critical segment of the aircraft’s design mission—a long endurance loiter. Low pitching moment coefficient allows for a weight savings from the use of smaller tails or other trimming devices. High lift-to-drag ratio during cruise saves fuel, allowing for extended loitering or lessens the requirement for onboard fuel. Gentle stall characteristics give the flight computer of the autonomous Martian aircraft better opportunity to recover from a stall. Low angle of attack during loiter will provide optimal downward orientation for the scientific instruments to survey the Martian landscape for indications of past liquid water. Finally, the high maximum lift coefficient will allow for a low sink rate on the landing approach as dictated by mission requirements.
The objectives were rank-ordered using the comparison matrix below and then the team assigned relative weights for use in evaluation of the airfoils.
Table #5 Airfoil Objectives Comparison Matrix
|
|
A |
B |
C |
D |
E |
F |
G |
Score |
|
A High loiter L/D |
--- |
1 |
1 |
1 |
1 |
1 |
½ |
5.5 |
|
B High cruise L/D |
0 |
--- |
1 |
½ |
½ |
1 |
0 |
3 |
|
C High max lift coeff. |
0 |
0 |
--- |
0 |
0 |
½ |
0 |
0.5 |
|
D Gentle stall |
0 |
½ |
1 |
--- |
½ |
1 |
0 |
3 |
|
E Low α during loiter |
0 |
½ |
1 |
½ |
--- |
1 |
½ |
2.5 |
|
F Low α during cruise |
0 |
0 |
½ |
0 |
0 |
--- |
0 |
0.5 |
|
G Low moment coeff. |
½ |
1 |
1 |
1 |
½ |
1 |
--- |
5 |
Table # 6
Assignment of Relative Weights to Airfoil Objectives:
|
Objective: |
Votes |
% |
|
High loiter L/D |
102 |
0.34 |
|
Low moment coefficient |
60 |
0.20 |
|
High cruise L/D |
48 |
0.16 |
|
Gentle stall |
42 |
0.14 |
|
Low α during loiter |
24 |
0.08 |
|
High max lift coefficient |
18 |
0.06 |
|
Low α during cruise |
6 |
0.02 |
The team then evaluated the airfoils using the scoring system and evaluation matrices that follow—comparing each airfoil against the others with respect to each selection objective.
Table # 7 Scoring system:
|
Best |
Better |
Average |
Poor |
Worst |
|
6 |
3 |
0 |
-3 |
-6 |
Table # 8 Evaluation Matrix for “Conventional” Airfoils:
|
Objective: |
Weight |
2415 |
4415 |
23015 |
65(216)-415 |
|
High loiter L/D |
0.34 |
0 |
3 |
-3 |
0 |
|
Low moment coeff. |
0.20 |
3 |
0 |
6 |
3 |
|
High cruise L/D |
0.16 |
3 |
0 |
0 |
6 |
|
Gentle stall |
0.14 |
-3 |
6 |
-6 |
3 |
|
Low α during loiter |
0.08 |
0 |
3 |
0 |
3 |
|
High max lift coeff. |
0.06 |
3 |
3 |
3 |
3 |
|
Low α during cruise |
0.02 |
0 |
3 |
0 |
3 |
|
Scores: |
|
0.84 |
2.34 |
-0.48 |
2.46 |
Table #9 Evaluation Matrix for “Low Reynolds Number” Airfoils:
|
Objective: |
Weight |
E214 |
E387 |
MB253515SM |
S4233 |
E205 |
S4180 |
SD7037 |
|
Loiter L/D |
0.34 |
6 |
0 |
-3 |
0 |
0 |
0 |
3 |
|
Moment coeff. |
0.20 |
0 |
0 |
0 |
0 |
0 |
0 |
0 |
|
Cruise L/D |
0.16 |
6 |
3 |
3 |
0 |
0 |
0 |
3 |
|
Gentle stall |
0.14 |
3 |
0 |
3 |
0 |
0 |
-3 |
0 |
|
Low loiter α |
0.08 |
3 |
3 |
3 |
0 |
3 |
0 |
0 |
|
Max lift coeff. |
0.06 |
6 |
6 |
0 |
3 |
0 |
6 |
3 |
|
Low cruise α |
0.02 |
6 |
3 |
0 |
0 |
3 |
0 |
0 |
|
Scores: |
|
4.14 |
1.14 |
0.12 |
0.18 |
0.3 |
-0.06 |
1.68 |
The design team decided to chose two “best” airfoils, one from among the four NACA airfoils and another from the seven low Re airfoils included in the selection process. The motivation for making this distinction is that the data for the NACA airfoils was obtained from tests with considerably higher Reynolds numbers than the Mars airplane will experience. This aforementioned experimental data collection condition results in an inaccurate reflection of the performance of these airfoils at extremely low Reynolds numbers. As a result of the high degree of uncertainty in the performance of the NACA airfoils, the team evaluated these four conventional airfoils separately from the seven low Re airfoils. In conducting the evaluation using the weighted objectives method, the NACA 65(216)-415 emerged as the best of the conventional airfoils while the Eppler E214 was the best of the low Reynolds number airfoils.
The evaluation process selected the NACA 65(216)-415 airfoil as the best of the conventional airfoils. This choice is reasonable because this airfoil exhibits average loiter lift-to-drag, a low moment coefficient, high cruise lift-to-drag, gentle stall characteristics, low angle of attack during loiter, and high maximum lift coefficient. This airfoil does not necessarily excel in meeting all the stated objectives; however, it is a decent overall airfoil—figuring at or above average in each objective.
The weighted objectives selection produced the Eppler E 214 as the best of the low Reynolds number airfoils. The E 214 choice largely results from its exceptional loiter and cruise lift-to-drag ratios as compared to the other six low Reynolds number airfoils. The E 214 outclasses the other airfoils considered in both these areas—standing out as the most efficient airfoil for operating at the low Reynolds numbers expected in the Martian atmosphere. In addition, the E 214 has an above average value for maximum lift coefficient, giving it the advantage of a slower rate of decent and a higher probability of a successful landing upon completion of the mission.
Ultimately the team concluded that selecting the E 214 for use on the Aries Alpha was the most prudent decision as more reliable experimental performance data is available for the E 214 at Reynolds numbers on the order of 100,000 than for the NACA 65(216)-415. The section profile, lift curve, and drag polar for the E 214 appear below.
Figure #12
Figure # 13 
3.2 Drag Build-Up
The drag
characteristics of the Aries Alpha
were computed using a parasite drag build-up involving each major component of
the aircraft’s configuration—wings, fuselage, and rocket engine nozzle. The drag of the wing was computed using an
interference factor of 1.0 since the Aries
Alpha is a mid-wing configuration.
The fineness ratio of the fuselage equals 4 and was used in computing
the wetted area and form factor en route to calculating the parasite drag
coefficient of the fuselage. The drag of
the rocket engine nozzle was approximated using the approach typically used for
nacelles along with dimensions of 0.094 meters length and 0.013 meters
diameter. The engine dimensions were
obtained from specifications provided by Professor Crossley for a Nitrogen
Tetroxide/MonoMethyl Hydrazine rocket thruster with specific impulse of 300
seconds and thrust of 111
Table # 10 Parasite Drag Build-Up
|
|
Loiter |
Cruise-Climb |
||
|
|
CDP |
fe [m2] |
CDP |
fe [m2] |
|
Wing |
0.0218 |
0.0414 |
0.0200 |
0.0380 |
|
Fuselage |
0.0085 |
0.0161 |
0.0078 |
0.0148 |
|
Nozzle |
0.000026 |
0.000049 |
0.000023 |
0.000044 |
|
Total |
0.0303 |
0.0575 |
0.0278 |
0.0528 |
3.3 Thrust
Requirement
Using the
parasite drag coefficients for the loiter and cruise-climb mission segments
along with average predicted atmospheric densities and flight velocities and
induced drag coefficients based upon the weight of the aircraft during each
segment, the team predicted the total drag forces expected during each of these
critical mission segments. During
cruise-climb, the Aries Alpha will
experience a drag force of 17.54
3.4 High Lift Devices
Following
the team’s decision, the Aries Alpha
does not have any high lift devices. The
rationale driving the choice to not utilize high lift devices stems from the
project focus on keeping weight to a minimum.
Focusing on achieving the smallest aircraft weight feasible, the team
did not include high lift devices in the sizing process—which ultimately
indicated that no additional lift devices are needed to remain under the
maximum sink rate constraint. However,
following detailed analysis of the aircraft’s takeoff sequence—in which it was
determined that the solid rocket booster originally chosen before sizing the
aircraft is does not burn long enough to accelerate the aircraft above its stall
velocity—the team believes that consideration should be given to flaps or other
high lift devices in further refinements to the design concept.
4. Stability and Control
4.1
There are many factors to consider when addressing the challenges of maintaining static and dynamic stability in flight. First and foremost are weight distribution, center of gravity, and the neutral point. It has been established that an aircraft, to be stable in steady level flight, must have the center of gravity placed forward of the neutral point according to Karl Nickel and Michael Wohlfahrt. For the purposes of this design class, the neutral point or aerodynamic centre was approximated to be at the quarter chord of the wing. The center of gravity placement carries an even greater importance for tailless aircraft, as they do not have extra control surfaces to trim flight. To achieve the longitudinal and pitch moment stability benefits of this convention, interior placement of system mass as well as total aircraft sizing must be taken into consideration. The location of structural mass about the aircraft can be tailored to lend itself to balancing with the interior mass, from such systems as fuel tanks and scientific instruments. In the process of affirming the decision to proceed with the tailless design configuration, fuel estimates and given payload values were compared to show feasibility of this stable aircraft design concept. Many of the internal systems are static with respect to their mass, such as avionics, landing gear, and the engine. However, fuel tanks which are full at take off are empty by landing, and from this the center of gravity is relocated. Since the fuel tanks are located near the center of the aircraft, the center of gravity will shift forward during flight toward the static mass systems. This is a clear benefit to the interior layout (Figure #11) as a negative nose pitching moment on the aircraft is easier to manage with control surfaces than a positive pitching moment. The calculation of the center of gravity, briefly outlined involved the centering of mass figures from the internal components based on volume and total mass, then algebraically calculating the moments and balancing them around a point. That length from the nose is 0.836 m, and is the location of the Aries Alpha center of gravity. The following Figure is a rough graphical representation of the major system weights whose respective masses were used in the center of gravity calculation. The gold and black ellipse designates the center of gravity location.
Figure #14
4.2 Trim Device
Considerations
During the design process, a number of configurations were considered in an attempt to optimize aircraft performance and stability in the Martian atmosphere. Chief among the considerations was the addition of a canard to the aircraft. The benefits of a canard design include but are not limited to; positive lift from these surfaces in trimmed flight, increased longitudinal stability, increase in potential control surface area (without the use of more wing planform area), avoidance of deep stall phenomenon, and aesthetic appeal. In brief, there are many valid benefits of this design. However the opportunity cost of this component was not equitable with the prospect of flying without it. The first reason to avoid a canard design was that the design concept from which this aircraft was based (the Darkstar UAV) did not include this feature. Further more, the sizing of the aircraft due to the addition of a canard presents the possibility of added costs and constraints. To avoid an increase in interference drag and to optimize lift distribution on both wing and canard, distancing the two is necessary. This involves a lengthening of the fuselage of the aircraft. Such alterations in proportions increases structural mass and generally, produce a larger sized aircraft. With a large amount of the cost and physical constraint for this mission profile coming from the launch vehicle itself, a design must minimize mass and spacecraft volume required for the transfer to the Mars operational environment. In comparison the additional cost of launch vehicle size increases and mass budget increases show themselves to be of much greater concern than the alternative solutions to stability issues.
4.3 Tailless
Stability Considerations
The methods used to maintain stability in flight beyond center of gravity placement included robust computer control programming and control surface sizing and configuration. With a mission of this magnitude and importance, the proper computer control coding is vital. To account for a changing environment and complex attitude adjustment, a proprietary code such as that utilized by the Northrop Grumman B-2 Spirit or the Darkstar UAV would be acquired and employed for this aircraft. Another focal point of stability and control concerns that arise with a tailless aircraft are the control surfaces themselves. To properly discern a method of control, a design database was compiled and analyzed for successful control surface configurations. Of the types considered were; the B-2 Spirit, Darkstar UAV, and Northrop Grumman XB-35. Following this process, a configuration similar to the B-2 Spirit was adopted, with similar control surface sizing, based upon percentage of the total wing area. Modeled after the B-2 Spirit, the control surface area was estimated as approximately 15% of the wing area for the Aries Alpha. Control in the pitch, roll, and yaw axes is obtained by the ruddervator and aileron devices. The pitch control and longitudinal stability, which is the most volatile for tailless aircraft, is governed by a combination of the ruddervators and the inboard flaps (Figure #10). Yaw stability and control are governed by the ruddervators themselves, as they split and deflect airflow both upwards and downwards, producing an induced drag effect. Roll control is maintained with the ailerons similar to standard tailed aircraft. Basically, all three axes require constant control surface deflection to keep the aircraft in trimmed flight. This factor must not be overlooked as it effects drag buildup and lift distribution on the wing planform area. To account for this in the performance of the Aries Alpha, a correction factor of 20% was added to the induced drag calculation. With this correction factor, a more conservative flight performance estimate was established.
There are many other methods to optimize stability and direct control for a tailless aircraft. For the purposes of center of gravity placement and compressibility drag decrement, wing sweep is an effective layout. The sweep of the wing comes at a price however. As the sweep increases, so does the spar length to accommodate the longer diagonal wing. Thus structural mass is increased, and the volume required to transport this craft in space is increased. According to Karl Nickel and Michael Wohlfahrt, the use of wing wash-out is a viable method of balancing pitching moments about the aircraft to sustain stability. This technique allows for a differing angle of incidence along the wing so that the wing-root stalls at a different time than the wing-tip. While this methodology is certainly useful, its incorporation into the aircraft design was unable to be completed. Nonetheless, it was investigated for the purpose of discussion and enhancing engineering understanding. Dihedral angles were considered an option for roll stability but were not applied do to the large amount of control surface area already applied to that axis of control. Yet another alternative technique of stability includes the use of positive moment coefficient airfoils. Most airfoils have a negative moment coefficient in regular flight and therefore lend themselves to instability in a tailless design where a balancing force from elevators is not present. While this is a most attractive feature for an aircraft like the Aries Alpha, it was not employed as airfoil selection centered on low Reynolds number performance. Considering that low Reynolds number airfoils are rare to begin with, a positive moment airfoil constraint might yield a trade-off of low density lift performance against moment coefficient which is a battle won by other methods as described previously.
5. Propulsion
5.1 Engine Size and
Thrust
In selecting the Aries Alpha’s engine, carpet plots and
sizing code generated the aircraft’s gross takeoff weight and thrust-to-weight
ratio. The product of these values yielded a target thrust output for the
aircraft engine of 49
5.2 Propellant
Consumption
Using re-arranged forms of the Breguet range and endurance equations, the fuel weight required for cruise-climb and loiter was determined in an iterative manner over small increments of each flight segment.
Table #1
|
Takeoff |
16.36 N |
|
Climb to 500 m |
21.06 N |
|
Cruise-climb (40 km) |
14.27 N |
|
Descent |
0 N (assumed no range credit) |
|
Loiter |
113.22 N |
|
Final Gliding Descent |
0 N (aircraft fuel exhausted) |
|
Total Fuel Weight |
165.52 N |
6. Performance
6.1 Rocket Assisted
Takeoff
The design team
decided a rocket assisted takeoff (RATO) would be necessary without the
provision of a prepared runway. In order to achieve a rapid acceleration to a
velocity at which the Aries Alpha
could commence wing borne flight, it was estimated that a thrust-to-weight
ratio approaching two would be needed. After researching solid rocket motors,
the Hypertek J315 was chosen as it was predicted to satisfy this requirement.
Since the solid rocket booster (SRB) has a mass of 1.8 kilograms, a mass of 2
kilograms was approximated in order to account for the mass of the attachment
mechanism, which will connect to the underside of the aircraft and later
jettison the SRB after its fuel is expended. The SRB has a max thrust of 566
This burn time discrepancy can be addressed by considering use of high lift devices and a larger SRB during subsequent refinements of the design concept. The addition of high lift devices and a change in the SRB would require a complete resizing of the aircraft as high lift devices would increase the empty weight and a larger SRB would increase takeoff / fuel weight. The team believes that taking these measures into account could provide a feasible solution to this issue.
6.2 Flight Profile
Figure #14
Table #12 Aircraft Performance
|
|
168.29 m/s |
|
Peak Altitude of Cruise-Climb |
1220.3 m |
|
Duration of Cruise-Climb |
3.96 min |
|
Average Loiter Velocity |
103.01 m/s |
|
Distance Traveled During Loiter |
278.1 km |
|
Sink Rate During Final Descent |
7.346 m/s |
Table #13
Aircraft Weight
|
Takeoff weight |
317.2 N |
|
Weight after takeoff and SRB jettison |
300.84 N |
|
Weight after climbing to 500m |
279.78 N |
|
Weight at end of cruise-climb |
265.51 N |
|
Weight at end of loiter |
152.29 N |
|
Empty weight |
152.29 N |
Figure #15
The Aries Alpha flies the design mission outlined in Figure #14,
achieving the performance numbers listed in Table #12, and experiencing weight
decreases as described in Table #13.
Figure #15 provides graphical verification that the loiter velocity
remains above the stall velocity (computed taking weight decreases into
account) throughout the duration of the loiter mission segment.
6.3 Descent and
Landing
After the aircraft has completed its 40 km cruise and 45 minute loiter through the Valles Marineris, its fuel expended, the aircraft will perform a gliding descent at a sink rate of 7.35 m/s to a height of 20 m above the canyon floor. At this point, the aircraft will be traveling with a forward velocity of 90.78 m/s, which is 5% above its empty weight stall speed of 86.46 m/s. Upon reaching the 20 m height, the aircraft will execute a flare maneuver while a lower panel of the aircraft blows off. Simultaneously, the instrument package is released and falls through the opening towards the Martian surface. Subsequently, small tanks filled with compressed gas inflate an airbag landing cushion system around the instrument package. The instrument package impacts the Martian surface with a vertical velocity component of 14.24 m/s, where it bounces and rolls to a safe stop, deflates the airbags and rights itself, and begins transmitting information via the Planned Mars Network. The decision to use the method of jettison and airbag landing of the instrument system was made to account for Mars’ rough, unpredictable terrain and the high forward velocity at landing. The design team reasoned that the probability of successfully landing the aircraft in a controlled fashion was small due to uneven Martian surface features and boulders. A crash—caused for instance by a wing clipping a boulder—at the approach velocity of 90.78 m/s, would likely render the aircraft’s communication antennae useless for data transmission and may destroy the entire aircraft.
As a result of these considerations, it was decided to eject the instrument package rather than attempt to land the aircraft. The airbag landing system used to preserve the instruments and transmitters for relaying scientific data is modeled after the Mars Pathfinder landing. This landing method allows for a more reliable method of landing the crucial instrument package intact. Rather than attempting to land and stop quickly or risk a massive impact with a surface boulder, the airbag system allows for the momentum of the instrument package to be dissipated gradually as it bounces and rolls along the floor of the Valles Marineris before coming to rest. After reaching its final resting place, the airbags deflate and the instrument unit rights itself to its designed data transmission orientation.
Figure # 16
Depiction of
instrument jettison and airbag landing at the conclusion of the flight.
7. Weights
7.1 Estimation
Methods
One of the most important steps in sizing the aircraft is the accurate prediction of the vehicle’s empty weight. To predict the wing weight, the team used Equation 8.1 from Brandt, et. al. which is typically used for prediction of jet aircraft wing weight.
Wwing = [0.04 S
(nmax)0.2 AR1.8 (1+λ)0.5] / [(t/c)0.7cos(ΛLE)]
In this process, a maximum load factor of 3 was assumed (equivalent to that for terrestrial transport aircraft). Realizing that the equation’s heavy dependence upon aspect ratio can cause significant overestimates of wing weight for high aspect ratio wings, the team decided to verify the accuracy of equation for predicting the empty weight of the Darkstar UAV which has an aspect ratio of 14.8. The Darkstar’s empty weight was predicted using the specifications for the Darkstar, again assuming a maximum load factor of 3, the installed engine factor for a jet transport (as the light aircraft factor was for propeller aircraft), the light aircraft factor for landing gear, the Darkstar’s instrument payload of 1000 pounds, and the miscellaneous weight factor for jet aircraft. The final predicted empty weight totaled 4650.7 pounds compared to the actual empty weight of 4500 pounds. Satisfied that the wing weight equation produced reliable results (in this case within 3.3% of the actual value), the team reasoned that no correction factor is needed to adjust the wing weight prediction for aspect ratios near 15.
The
fuselage weight was computed using the estimate for wetted area in conjunction
with the light aircraft fuselage factor (converted to units of
To estimate
the engine weight, the team obtained a weight estimate for a 111N thrust
NTO/MMH liquid rocket thruster from Professor Crossley. Converting the weight from pounds measured on
Earth to
Components
of the miscellaneous weight include the avionics, scientific instrumentation,
and other unaccounted for aircraft components (i.e. wires, fasteners,
etc.). Avionics mass is 10 kg, yielding
a weight of 37.2
7.2 Weight Breakdown
Table #14 Aircraft
Weight Breakdown [
|
Wing |
27.9746 |
|
|
Fuselage |
29.0768 |
|
|
Total Structure |
|
57.0514 |
|
Engine (uninstalled) |
|
2.9499 |
|
Landing System |
18.6000 |
|
|
Miscellaneous |
73.0452 |
|
|
Total Systems |
|
91.6452 |
|
Aircraft Empty Weight |
|
151.6466 |
III. Spacecraft
1. Launch Vehicle
1.1
Launch Vehicle Database
Table #15
|
|
Atlas III-B |
Atlas V (521) |
Atlas V (551) |
Delta II (7925-10) |
Delta III |
Delta IV (medium) |
|
Max Payload Weight (kg) |
4500 |
6000 |
8200 |
1799 |
3810 |
3900 |
|
Alt/radii to Apogee (km) |
185 |
185 |
185 |
185 |
185 |
185 |
|
Alt/radii to Perigee (km) |
35786 |
35786 |
35786 |
35786 |
35786 |
35786 |
|
Max Length of Payload Env. (m) |
|
|
|
4.1 |
8.9 |
11.4 |
|
Diameter of Payload Envelope (m) |
3 |
4.57 |
4.57 |
3 |
4 |
4 |
|
Tot Launch Vehicle Mass (kg) |
225392 |
418725 |
541195 |
231670 |
301454 |
|
|
Launch Reliability |
50 launches |
|
|
100% after '97 |
|
100% |
|
Cost (in million dollars) |
98 |
95 |
110 |
52 |
83 |
90 |
|
|
|
|
|
|
|
|
|
|
Titan II |
Titan IV |
Ariane 4 |
Ariane 5 |
Proton K |
Zenit 2 |
|
Max Payload Weight (kg) |
|
6350 |
2175 |
6800 |
4815 |
6000 |
|
Alt/radii to Apogee (km) |
185 |
185 |
185 |
185 |
185 |
185 |
|
Alt/radii to Perigee (km) |
35786 |
35786 |
35786 |
35786 |
35786 |
35786 |
|
Max Length of Payload Env. (m) |
7.6 |
15.7 |
7.12 |
11.5 |
7.8 |
13.7 |
|
Diameter of Payload Envelope (m) |
3.1 |
4.3 |
5 |
5.4 |
3.8 |
|
|
Tot Launch Vehicle Mass (kg) |
|
|
|
737000 |
691272 |
|
|
Launch Reliability |
100% |
91% |
1 |
90% |
96% |
|
|
Cost (in million dollars) |
35 |
89 |
85 |
120 |
|
79 |
1.2 Launch Vehicle
Selection
The method for the determination of an appropriate launch vehicle was as follows. First, the team calculated the total delta-v required to boost the spacecraft from the elliptic Geosynchronous Transfer Orbit (GTO), onto an elliptic, heliocentric Hohmann transfer to Mars, and finally enter Martian orbit and perform an entry burn. The team wrote a MATLAB function which would predict—based upon an input initial launch mass, the propellant required for a particular specific impulse spacecraft engine, and historical planetary spacecraft mass breakdown percentages—the aircraft mass a particular launch vehicle is capable of launching for this mission. Then, inputting the mass to GTO for various launch vehicles and the specific impulse of the spacecraft engine into the MATLAB function, the aircraft payload that would actually be sent to Mars without cutting into the dry mass margin was computed for several launch vehicles selected from the database.
1.3 Selected Launch
Vehicle
Based on mass capable of being boosted to GTO and the capacity for the Aries Alpha’s mass and dimensions, a launch vehicle was selected: the Atlas V (551). This launch vehicle has a total mass of 541,195 kilograms, maximum payload mass to GTO of 8,200 kilograms, five solid rocket boosters, and a usable payload diameter of 4.57 meters. The estimated launch price of the Atlas V (551) is 110 million dollars.
Figure #18 Figure #19
2. Spacecraft Description
2.1 General
Description
The spacecraft to be used during interplanetary flight will be a custom designed vehicle. The primary purpose of the spacecraft is to provide the necessary propulsion to send the entry vehicle containing the aircraft from GTO, onto the heliocentric transfer orbit to Mars, enter Martian orbit, and finally decelerate enough to enter the Martian atmosphere. The team selected the Pratt and Whitney RL-10B-2 liquid bi-propellant rocket to meet this fundamental requirement. With a specific impulse of 465.5 seconds, this engine was one of few existing engines capable of performing this flight with the selected launch vehicle and mass of the payload. Furthermore, the RL-10 series of engines has, for forty years, been the most reliable engine line available. In addition, the entry vehicle must provide thermal protection for the aircraft payload and a means of decelerating to a safe landing descent velocity. This will be accomplished using an aeroshield/parachute/airbag entry and landing sequence similar to that of Mars Pathfinder. Upon successful landing, the spacecraft must right itself and open into the takeoff configuration for deployment of the aircraft. The airbags will deflate in an order that helps to turn the cylindrical entry vehicle to its side where its doors will open end sections detach, allowing the airplane launch ramp to raise and the wings to unfold prior to takeoff.
2.2 General
Arrangement and Layout
Located immediately in front of the engine is the liquid hydrogen tank. Although the mass of liquid hydrogen used will be much less than the mass of oxygen used, liquid hydrogen is much less dense than liquid oxygen. Therefore, the liquid hydrogen tank is much larger than the liquid oxygen tank. The liquid oxygen tank rests directly in front of the liquid hydrogen tank, and behind the heat shield. The heat shield is the bottom of the landing craft. The diameter of the heat shield is much larger than the diameters of the engine, fuel/oxidizer tanks, or length of the aircraft. This increased size is necessary to protect the landing craft from the intense frictional heating encountered as it enters the Martian atmosphere. During the flight from Earth to Mars, the landing craft will also be protected from micro-meteoroid impacts by large outer panels on the spacecraft. These panels will also aid in thermal protection and aerodynamics during entry and will be jettisoned along with the heat shield as the parachute deploys.
2.3 Spacecraft Model
Figure #20
Figure #21
2.4 Spacecraft Launch
and Payload Mass
The total spacecraft mass at
launch is 8200 kg including the spacecraft with its fully fueled mass of 7530.9
kg and the spacecraft adapter mass of 669.1 kg.
This launch mass combined with the RL-10-B upper stage engine Isp of
465.5 seconds allows for an aircraft (payload) of 82.8 kg to be delivered to
the Martian surface. This figure retains
a spacecraft dry mass margin of 405.34 kg to allow for uncertainties at the
conceptual design stage. The Aries Alpha has a takeoff mass of 85.3
kg, however the additional 2.5 kg could be obtained by cutting into the mass
margin as design refinements continue.
This would result in using about 0.6% of the margin to allow for the
difference in takeoff mass of the aircraft and payload capability of the launch
vehicle and upper stage engine combination.
3. Astrodynamics
3.1 Earth to Mars
Trajectory
The Earth to Mars trajectory was designed using a patched conic approximation. This method assumes that Earth and Mars both follow circular orbits and that their orbits are co-planar. As a result, the trajectory design does not consider the changing velocity of the planets as the radius of the orbit changes, nor does it include a plane change maneuver to compensate for the 1.85 degree inclination difference with respect to the ecliptic plane.
The launch vehicle will serve to
boost the spacecraft to GTO. From GTO,
the upper stage spacecraft engine will ignite to accelerate the craft through a
delta V of 1.173 km/s in order to leave GTO at perigee and enter a hyperbolic
departure orbit, leaving Earth’s sphere of influence. Once it exits Earth’s sphere of influence, at
a distance of approximately 924000 km, the spacecraft will be on a
heliocentric, elliptical Hohmann transfer orbit with aphelion coinciding with
the position of Mars on the arrival date.
Upon arriving at aphelion, the spacecraft engine must provide another
delta V of 2.650 km/s in order to match the speed of Mars in that point of the
planet’s orbit. Next, the engine must
decelerate the spacecraft to enter Martian orbit at an altitude of 500 km—a
maneuver that requires a delta V of -1.373 km/s.
Figure #22
3.2 Atmosphere Entry
Approximation Orbit
The final role of the spacecraft engine is a small delta V to leave the 500 km altitude circular orbit about Mars and enter the Martian atmosphere. The entry orbit has a periapsis of 80 km to approximate the atmospheric entry trajectory. This final delta V is
-0.096 km/s, and completes the mission of the upper stage
engine. After completion of this
de-orbit burn, the spacecraft engine will be jettisoned and the entry vehicle
will begin its entry into the Martian atmosphere with the newly uncovered heat
shield pointed in the direction of the vehicle’s velocity.
3.3 Required Delta V
The total delta V (magnitude)
required from the spacecraft engine is 5.292 km/s for the entire mission from
GTO to the entry burn. This figure,
along with spacecraft engine specific impulse, determines the fuel mass
required for the mission. The specific
impulse and its impact on the fuel mass required for the mission is a driving
factor in the selection of the spacecraft engine.
4. Propulsion
4.1 Spacecraft Engine
Selection
Estimates
and calculations determined a spacecraft propulsion system with a specific
impulse above 460 seconds would be needed for this mission. After reviewing several possible engines,
only a few had specific impulses which could meet the requirements of this
mission. Two engines considered were
made in
4.2 The Upper Stage
Engine
The RL-10B-2 is a cryogenic, second-stage engine that uses liquid hydrogen (LH2) as its propellant and liquid oxygen (LOX) for an oxidizer. It has a total mass of 277 kg, a length of 4.14 meters, and a diameter of 2.21 meters. The diameter of the base engine bell is 2.1 meters. Along with a specific impulse of 465.5 seconds, the RL-10B-2 produces a total thrust of 24,750 lb. The mixture ratio for the B-2 model is 5.85 parts liquid oxygen to one part liquid hydrogen.
The original RL-10 was introduced nearly forty years ago. It was the first LOX/LH2 rocket to be operated in space. Since its initial use, it has become the most reliable upper-stage engine for space travel. The RL-10B-2 is simply an upgrade of that original, reliable design. This model uses the same basic engine and turbopumps as the original RL-10. It also uses flight-proven off-the-shelf components for its attitude control system (ACS).
In addition to the proven design basis and high specific impulse, the RL-10B-2 has many performance-enhancing features. The engine has an extendable exit cone that can increase the specific impulse and payload capability of the rocket. Electromagnetic actuators, instead of hydraulic pumps, control thrust gimbals on the rocket. This change increases the reliability of the gimbal system while simultaneously reducing engine weight and cost. An extra helium bottle for tank repressurization can also be added to the RL-10B-2 if more than one restart is needed during the flight. This feature makes the engine well suited for a flight to mars, where several large velocity changing burns will be needed.
4.3 Propellant
Consumption
The amount of propellant required to provide the delta Vs needed during the mission’s four burns is listed below. This allows a direct comparison of the propellant consumed in each segment of the trajectory maneuvering.
Table #16 Propellant Budget
|
Burn |
Delta-V (km/s) |
Propellant Mass |
|
Leave Earth GTO |
1.173 |
1705.97 kg |
|
Mars Arrival |
2.650 |
2564.56 kg |
|
Enter Mars Orbit |
-1.373 |
846.63 kg |
|
Atmospheric Entry |
-0.096 |
50.21 kg |
The useable propellant mass for the spacecraft is 5167.37 kg. Since some propellant will get “trapped” in the tanks and never consumed, an extra two percent (103.35 kg) will be added to the propellant mass. In addition to the extra two percent, 100 kg of reserve propellant will be added. With the reserve and trapped propellant additions, the total propellant mass for the spacecraft will be 5370.72 kg. Since the RL-10B-2 has a mixture ratio of 5.85:1, in favor of oxidizer, the total LH2 mass will be 784.05 kg. The mass of LOX will be 4586.67 kg. LH2 has a density of 70.8 kg/m3 and LOX has a density of 1140 kg/m3. Using the given density values, the tank volumes* for each of these propellants are 2925.48 gallons for LH2 and 1062.87 gallons for LOX.
*Note: Even though the mass of LOX needed is much greater
than the mass requirement of LH2, LOX is much denser than LH2 and the volume of
LH2 needed will be greater than the volume of LOX.
5. Spacecraft Mars Entry Sequence
5.1 Landing
Once the spacecraft reaches Martian orbit, it must enter the Martian atmosphere as the final stage of its flight. The entire landing sequence is similar to that of the Mars Pathfinder mission. At approximately 10 km altitude, a parachute will deploy and slow the craft to 60 m/s. Having served their purpose, the heat shield and protective outer panels will be jettisoned from the entry vehicle. The landing craft will use a radar altimeter to determine its altitude during descent. At an altitude of approximately 300 m the airbags will deploy encasing the landing craft, and solid rocket boosters will fire slowing the craft to almost 0 m/s at an altitude of nearly 20 m. The landing craft will then be released from the parachute to fall freely toward the Martian surface, where it will bounce on its airbag landing system and come to rest. Subsequently, the airbags will be deflated asymmetrically to orient the landing craft in the correct orientation. Orienting the craft will be aided by the weight distribution of the landing craft. All the instruments will be placed on the bottom side of the landing craft, underneath the aircraft and its launch ramp platform. This will create a tendency in the landing craft to roll so that side is on the bottom. In order to balance the craft’s center of gravity for rocket burns in transit to Mars, the packing of the parachutes will be placed opposite that of the instruments.
Figure #23
5.2 Preparation for
Aircraft Takeoff
After the airbags have deflated and the landing craft is on its side, the hydraulic doors will open, and will put the landing craft in the exact orientation for launch of the aircraft. The launch ramp will raise to an angle of 60°, where the aircraft will then be launched using solid rocket boosters. The solid rocket booster, combined with the aircraft engine, will provide the necessary thrust to accelerate the aircraft to a velocity greater than the aircrafts stall speed. The aircraft will then begin a cruise climb under the power of its own engine.
Figures #24 through 28
Figures #24-28 depict the opening of the spacecraft and
preparations for aircraft takeoff.
6. Mass
6.1 Estimation
To estimate the mass allocation of the spacecraft, an examination was conducted of previous missions. Looking at past missions exposed trends in the allocation of the available mass. These trends were used to estimate the mass allocation of this Mars mission. First, the spacecraft adapter mass was determined using a linear equation, developed from historical trends, with total mass to GTO as the independent variable. This equation returned a value of 669.1 kg for the spacecraft adapter. Next, using the calculated ΔV needed for transfer from Earth to Mars, the propellant mass was calculated. It was determined that, with margins for trapped and reserve propellant, 5370.7 kg of propellant would be needed for the space flight. The adapter and propellant masses were then subtracted from the total mass to GTO of the launch vehicle. This value represents the spacecraft dry mass.
A mass margin of 35% was used to account for errors at this early stage in the design process. With this percent, the margin mass is 405.34 kg. Since the propellant fraction for a liquid rocket engine is .9, the approximate mass of the engine itself is 596.74 kg. Both the margin mass and engine mass were subtracted from the spacecraft dry mass to find the allocated dry mass. The atmosphere entry system mass estimate was 45% of the allocated dry mass. Subtracting the entry system mass from the allocated dry mass, 636.97 kg was left for the major spacecraft subsystems. Finally, historical trends were again used to determine what percentage of the remaining allocated dry mass went to each subsystem.
Table #17 Spacecraft Dry Mass
|
Atlas V 551 8200
kg to GTO Space craft ISP of 465.5 sec |
||
|
subsystem |
mass budget allocation |
mass (kg) |
|
ACS |
10% |
63.70 |
|
C&DH |
7% |
44.59 |
|
Cabling |
8% |
50.96 |
|
Communications |
7% |
44.59 |
|
Payload |
13% |
82.80 |
|
Power |
22% |
140.13 |
|
Structure |
30% |
191.09 |
|
Thermal |
3% |
19.11 |
|
Allocated |
636.97 |
|
|
Atmosphere Entry System |
521.15 |
|
|
Propulsion |
596.74 |
|
|
Margin |
405.34 |
|
|
Total |
2160.20 |
|
III. System Cost
1. System Cost Discussion
No matter what the mission, operational cost is always a design and program driver. In this mission, cost was not defined by a stated value; however, a goal of this mission was to keep costs to a minimum.
The funding for this program must be sufficient to furnish the research and development, materials, engineering, production, deployment, and operation of systems costs for it to be a reality. In understanding this, a piecewise approach was applied to those areas for the different segments of the mission. The following table outlines major mission cost estimate contribution areas.
Table #18
|
Component Description |
Cost (in FY 1999 USD, Millions) |
|
Aircraft Systems Development and Construction Cost |
$22 M |
|
Launch Vehicle Cost Lockheed Martin Atlas V 551 |
$110 M |
|
Spacecraft Cost (Interplanetary Stage) |
$654.911 M |
|
Program Systems Operational Cost (1 Year) |
$22.7 M |
|
Fuel Cost (for Aircraft) |
$0.00248 M |
|
Total |