Executive Summary - 10 Pages Maximum (Separately Bound)

CROMAGnAM

Crew for Return of Multiple Atmosphere Gravity Assist Mission

A&AE 450

Fall 2005

Brad Crosson

Lloyd Droppers

Rob Fink

Mike Grant

Missy Hartwell

Greg Henning

Thomas Jennings

Tim Kite

Robert MacDermott

Michael Moessner

Tony Sanders

December 5th 2005

Executive Summary- Lloyd Droppers

Purpose and Goals

The purpose of this report is to evaluate the use of an aero-gravity-assist (AGA) at Mars and Venus, using methods simple enough to allow for multiple iterations in one semester of work and in-depth enough to determine the technical feasibility of the mission. The goal of the mission was to utilized aero-gravity-assist at Mars and Venus in order to sample planetary atmospheres and return the sample to Earth. The hope was that by utilizing aerodynamic forces a lighter and cheaper vehicle can be used as compared to a pure propulsion mission. The vehicle performance was evaluated using planetary trajectory entrance and exit conditions at Earth, Mars, and Venus provided by Tracy Smith given in the appendix. A vehicle that converged on the conditions was not found during the work performed this semester. Evaluated from a structure, propulsion, or thermal protection view the vehicle performed satisfactorily, however aerodynamic performance was not adequate, due to the low lift to drag ratio. Work done at the end of the semester led to a potentially improved lift to drag ratio, and indicated that the mission would be feasible with this improved aerodynamic performance.

Prior Work and Requirements

Previous interplanetary missions have used both gravity assist maneuvers and aerodynamic force to improve mission design. Gravity assists were used by Cassini-Huygens, MESSENGER, Galileo, and many other missions and has become a standard procedure to adjust a probe’s trajectory. Aerodynamic forces were used every shuttle mission and have been used on probes such as the Mars Exploration Rovers and Mars Odyssey missions. These mission use aerodynamic forces to decelerate the vehicle however, this was significantly different than the proposed mission. Aerodynamic lift would be used in conjunction with a gravity assist to achieve an increased effectiveness in the gravity assist.

While no mission to date has used an AGA, analysis has been done on AGA missions. Theses analysis generally assumed fixed lift to drag ratios and did not provide analysis of a thermal protection system (TPS), the effect of aeroshell volume and mass, and various other design factors. The primary task laid out by the Request for Proposal was to “provide an improved preliminary design for an AGA mission.” The interplanetary trajectories were calculated by Tracey Smith, a Purdue graduate student, using a simplified model for the vehicle and aerodynamics.

The design was initially required to carry a spacecraft similar to Galileo with a dry mass of 1300 kg and a V of 1400 m/s to perform interplanetary orbital maneuvers. The requirements were later loosened to allow for the dry mass to be at the discretion of system to include sufficient instrumentation and mechanisms to allow for the capturing of atmospheric samples, and the V reduced to 300 m/s for atmospheric maneuvering, and small trajectory corrections. There was a requirement to evaluate an ablative TPS, and design the “vehicle aeroshell shape, TPS, planet-encounter orbital dynamics, structure, and propulsion system.”

Assumptions

Due to the time constraints posed by the semester long preliminary design a variety of assumptions were made. The preliminary systems level assumption was that of disregarding producibility concerns.Ultra High Temperature Ceramics (UHTC) was assumed to be available with a leading edge radius as small as 1mm. Structurally the component mounting and joints was disregarded instead using and empirical scaling factor. The vehicle was assumed to have instantaneous response time for angle of attack and bank angle profile. The aerodynamic lift was assumed to be pure Newtonian and drag was assumed to consist of Newtonian, skin friction, and viscous interaction drag. These are the main assumptions of concern; each section of the report can be consulted for discipline specific assumptions.

Vehicle Design

In order to expedite the process of designing the vehicle certain initial shapes were decided upon. The vehicle was established as a flat sided wedge with ablative TPS on the windward half and a UHTC nose. The vehicle propulsion system was determined to be a pressure fed MMH/NTO for both the reaction control systems and the main engines. The vehicle was determined to have a graphite composite honeycomb monocoque structure. Initially the vehicle was relatively stubby, eventually a long thin wedge was determined to provide the best performance, and the vehicle iterations can be seen in Table 1.

Table 1: Vehicle Iteration History

The final iteration gave the best performance of the full iteration and a 3-view drawing can be see in Figure 1, and the mass information can be seen in Table 1. Each subsystem’s information can be seen in the next section.

Figure 1: Final Iteration 3-view drawing

Subsystem Performance

The vehicle that was designed was named SPEAR. This stands for Spacecraft for the Purpose of Exo-Atmosphere Rendezvous.

Systems- M. Moessner

The components of SPEAR met the minimum spacecraft requirements. It has scientific components, propulsion, data communication, avionics, instrumentation, and power components. These components were placed in the spacecraft to provide a stable CG location and to fit in the structural body.

Aerodynamics- M. Hartwell

In order to perform this mission, a hypersonic vehicle with a high L/D was desirable. A wedge with a hemicylindrical nose and a flap attached to the back for trim control and stability concerns was chosen as the final vehicle design based on its flat plate characteristics when flown at an angle of attack greater than the included angle. The Newtonian Hypersonic theory was implemented evaluate the vehicle performance. The vehicle shape as well as the viscous interaction drag was the main factor of the L/D and performance. See figure of Cd vs. Altitude. A perfectly blunt flat plate with an infinitesimally small nose radius has an in-viscid L/D approaching infinity; however, this was both impractical for internal volume as well as manufacturing capabilities. Viscous interaction can account for up to 70% of the total drag at high altitudes (for Mars) and 30% at lower altitudes. Therefore, it is very unreasonable to ignore viscous effects, especially when performing an aero gravity assist mission around two planets.

Aeroheating- B. Crosson

SPEAR was able to successfully traverse through the atmospheres of Mars, Venus, and Earth without incurring excessive heating that would result in mission failure. The end piece of the spacecraft shadowed the engines from the effects of heating and also allowed for a more manageable flap size. The nose was made up of ultra high temperature ceramic, UHTC, which would take on the hottest temperatures which occurred at the stagnation point. The UHTC was found to adequately withstand the heating effects during the AGA maneuvers and reentry to Earth.

TPS- R.MacDermott

The Thermal Protection System (TPS) was meant to protect the space craft against the heat associated with reentry into atmospheres at high velocity. Using the Sandia One-Dimensional Direct and Indirect Thermal code developed by Sandia National Laboratories for use with the Apollo missions, a system comprising an ablator and an insulator was developed to provide adequate thermal protection. The ablator sublimes at a given temperature, thus removing energy from the system, while the insulator keeps the graphite-epoxy composite skin and the internal components inside their optimal temperature range.

Structures- L. Droppers

The structural system is meant to support the vehicle components though out the flight for various loads. The structure was analyzed for a static launch acceleration loading and a aerodynamic deceleration loading. A simplified shell model was use analyze the graphite epoxy composite honeycomb structure. For the final vehicle and all previous iterations the worst case loading was the buckling from static launch acceleration loading, shown in Figure 3. To allow a 1.4 factor of safety as required by MSFC-HDBK-505, a faceplate thickness of .8mm and a 1.3mm honeycomb thickness was used. The structural mass percent was a reasonably low 14% of dry vehicle mass, the structural weight remained reasonable as long the vehicle stay reasonable slender, roughly under a 3 to 1 length to width ratio.

Propulsion- T. Jennings

Propulsion will be included in a further revision of this report due to extenuating circumstances.

Mars Trajectory- M. Grant

The capability of the vehicle to perform the turn at Mars is influenced by the L/D performance of the vehicle throughout the trajectory. Various parameters influence the L/D performance that is capable at Mars. First, the nose radius must be reduced to smallest values attainable in order to provide higher L/D performance. Also, the viscous drag interaction is detrimental to the L/D performance of the vehicle. The vehicle must either be long or fly at low altitudes in order to minimize the impact of viscous drag interaction. The vehicle dimensions were reduced in order to increase the ballistic capability of the vehicle to overcome drag. This allowed the vehicle to fly at lower altitudes to minimize viscous drag interaction. However, higher angles of attack must be used with more compact and more massive vehicles. When the vehicle length was increased, the altitude profile was increased and smaller angles of attack were used throughout the trajectory. This provided the best L/D performance (see figure below). However, the optimal vehicle length and lifting surface dimensions must be further analyzed in order to improve L/D performance. The error in exit state for this best case is quite small (with a small underspeed) and can be found in the table below. Thus, the mission can be completed with minor gains in L/D performance.

Table 2: Best Scenario – Long Vehicle (10 m)

Venus Trajectory- G. Henning

The trajectory design for the Venus flyby took on many forms before the design was finalized. The initial and final states outside the atmosphere of Venus were given in terms of altitude, flight path angle, and velocity. Also given were the V∞ vectors for the approach and departure hyperbolas. From this information, it was possible to design a path from start to finish. The vehicle was controlled through the trajectory by varying its AOA and Bank angle. These were varied using different types of control schemes such as PID controller and simple if-then statements. After many iterations, it was determined that a simple open-loop system in which only a few maneuvers were performed at specified times in the flight was the most efficient method. The specified final conditions were all met within a small window of error, with the velocity error being the maximum at about 4% under the desired value. The main conclusion drawn from the trajectory point of view was either that a better L/D needs to be achieved through vehicle improvement, or that a more accurate aerodynamic model should be employed before designing the interplanetary portions of the flight. Assuming fixed CL and CD isn’t a reliable way to approximate the AGA mission.

Table 3:VenusExitState

Mars MDO- A. Sanders

The Mars atmosphere entry optimization took into major affect aerodynamic properties and Martian atmosphere affects on the aerodynamic properties. At first glance, it might seem that the Martian atmosphere would provide little hindrance to the missions’ success because the density of the atmosphere is so low. The problem that was run into was the very large viscous interaction effects that act on a vehicle in a low density situation. These viscous interaction effects greatly increase the total drag on the vehicle and have a very large effect on the velocity loss of the vehicle in the Martian atmosphere. In light of this fact, viscous interaction effects are a large role in the success of the aero-gravity-assist maneuver and must be taken into effect.

Venus MDO-R. Fink

Initially, optimizing the mission at the Venus stage could be narrowed down to the particular focus of the operating vessel, SPEAR, since this was within the scope of the primary analysis tools. There were many considerations in optimizing SPEAR to begin with, such as its total mass, ability to ablate heat, aerodynamic performance and capacity to satisfy trajectory requirements, among possible others. Over time, the prevalent concern that moved to the forefront of design optimization was SPEAR’s aerodynamic performance as it flew a trajectory through the Venusian atmosphere, since it needed a high enough L/D to do so. The particular focus of this element of the project was to try and change the vehicle dimensions to produce such a result. This was done by selectively altering particular dimensions, one at a time, and seeing if there was a better boundary or range near the given specs to get a better L/D in atmospheric flight. Ultimately, even though SPEAR could be conserved in size to produce less drag, the L/D desired could not be achieved.

Earth MDO – Tim Kite

The Earth re-entry stage of the design optimization process involved the determination of an optimized re-entry trajectory and control scheme which would allow SPEAR to re-enter the atmosphere given certain limiting parameters, such as maximum loading, stable angles of attack, and heating limitations, and without requiring changes to the spacecraft design that would negatively impact the main mission. This was done as part of the overall iterative process, taking sizing and stability information and returning structural loading and atmospheric heating data during each iterative cycle. After optimizing the trajectory and controller, the result was a successful re-entry with a decently large landing area.The hypersonic portion of the flight can be seen in Figure 5.

Figure 5: Landing Area

Conclusion

The vehicle as designed was unable to meet the requirements of the RFP. There are two main factors keeping the vehicle from meeting its objectives, aerodynamics and lack of including the interplanetary trajectory in the iterations. It is possible that if the interplanetary trajectory was included the vehicle as designed could meet a modified trajectory, but as this was not within the scope of this design.Future work would have to be done on this regard. Another possible method of meeting the requirement would be to modify the vehicle to improve aerodynamic performance as discussed in the Mars Trajectory section, and in further detailed in the main body of the report. The mission objectives could also be met by supplemental propulsion to provide the delta-V required to achieve the proper planetary exit flight conditions. The main understanding gathered by this investigation is that, with the vehicle analyzed, the mission was feasible with the caveat of improving aerodynamic performance, either through better understanding of the environment and viscous interaction drag or though improved aerodynamic design.

Appendix A

Appendix B

Interplanetary Trajectory information – Tracey Smith

Initial Conditions Results

Both of the following trajectories are launched from Earth on Feb. 20, 2016 with a launch V of 3.5 km/s. The maximum L/D ratios used in STOUR is 3 for each of the atmospheric fly-throughs; however, the actual value required to fly the trajectory will be slightly higher than 3.

Trajectory 1: Low V∞

Earth Departure

Departure Date: Feb. 20, 2016

V∞+ = 1.5766807193x + 2.9938697189y + 0.893105520z

Planet departure conditions

R_dep = 7278.13 km

V_dep = 11.0356 km/s

_dep = 0.151058 rad

Mars

Arrival Date: July 19, 2016

V∞- = 7.8772231541x + 1.0592589199y + 0.237393750z

V∞+ = 2.2816745842x - 3.1428622660y - 4.072837473z

Planet approach conditions

R_initial = 3847 km

V_initial = 9.24634200946673 km/s

_initial = -0.44388870203440 rad

Planet departure conditions

R_final = 3847.022246316240 km

V_final = 7.34421118271270 km/s

_final = 0.42406474986035 rad

Venus

Arrival Date: Jan. 25, 2017

V∞- = -2.5372107523x + 4.6309662682y - 3.1250119833z

V∞+ = -0.6142233463x + 2.6595382310y + 4.3204364298z

Planet approach conditions

R_initial = 6501.8 km

V_initial = 11.72935885010359 km/s

_initial = -0.26780287776034 rad

Planet departure conditions

R_final = 6501.327979712902 km

V_final = 11.22699679457943 km/s

_final = 0.26072646132083 rad

Earth Arrival

Arrival Date: July 3, 2017

V∞+ = 2.53511877461x - 0.1965298704y – 1.6167828300z

Planet arrival conditions

R_arr = 7278.13 km

V_arr = 10.8910 km/s

_arr = -0.1494207 rad