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 |
||