Orbital Space Plane

O R I O N

Orbital Space Plane

EXECUTIVE SUMMARY

12.08.03

Purdue University

AAE 450

BOOST VEHICLE DESIGN AND VEHICLE PROPULSION Josh Jung

TRAJECTORY/PERFORMANCE ANALYSIS Jim Palmer

AEROTHERMODYNAMICS Brian Schoening

STRUCTURES AND LOADS Jenny Drake

SYSTEMS Justin Bailey

The orbital space plane presented in this report (ORION) is the last in a series of iterations designed to optimize performance in five general disciplines: boost vehicle design & vehicle propulsion, trajectory/performance analysis, aerothermodynamics, structures & loads, and systems. The result is a design that maximizes the method used to the extent that its own limitations allow.

ORION is the middle of an idea that started when man first left Earth’s orbit. It pushes the limit on prospective technology, surpassing all requirements asked for in the RFP.

The first mission of ORION is to carry 7 astronauts to and from the Space Station. In an emergency, the astronauts will be able to reach definitive medical care within 24 hours. The crew does not have to wear suits while in the vehicle. It will be launched aboard a Delta IV-H and carry a payload of over 6000 kg.

The second mission of ORION is to serve as a military space plane flying reconnaissance missions and providing surveillance capabilities. Part of the MSP mission is to be able to perform an aero-assist maneuver to change the inclination of the space plane. To accomplish this a larger L/D is required, thus the aerodynamic design was in part concentrated on designing a vehicle with optimal aerodynamic characteristics.

However, it was determined at this stage in the design that converting ORION to an MSP is not feasible. The propulsion requirement to achieve enough ΔV to change orbital planes was too massive—including a reduction in propellant due to aero assist. Thus, the focus of this analysis will be on the NASA mission.

1.0  BOOST VEHICLE DESIGN AND VEHICLE PROPULSION

Boost Vehicle:

ORION must be boosted into ISS orbit (407 km altitude, 51.6° inclination) to satisfy the mission requirements. The Delta IV-Heavy rocket supplemented with a Centaur III as the second stage was selected to lift the 22757 kg ORION into the specified orbit. This launch vehicle consists of a core 1st stage, one boosters (similar to core), and one Centaur III 2nd stage. Each stage burns liquid oxygen and liquid hydrogen. Several important parameters for the launch vehicle are shown in Table 1-1 below.

Table 1-1: Important launch vehicle parameters.

Stage / Booster / Stage 1 / Stage 2
Designation / Delta IV Core / Centaur III
Diameter / [m] / 5.10 / 3.05
Length / [m] / 40.8 / 11.68
Thrust (vac) / [kN] / 3315 / 198.4
Isp (vac) / [s] / 420 / 450.5
m0 / [kg] / 226360 / 22960
tb / [s] / 248 / 464

The vehicle parameters were then input into a trajectory code, that employs Vinh’s equations for a rotating earth to calculate several parameters along the boost trajectory. These include: flight path angle, heading angle, latitude, longitude, thrust, mass, acceleration, velocity, altitude, and stagnation temperature. Impact sites of the spent stages were calculated and an analysis on gimbal control to counteract the effect of 99 percentile winds was also performed. A plot of the ground track is show in Figure 1-1.

Figure 1-1: Ground track and separated stage impact sites.

ORION Propulsion:

ORION consists of two joined systems: the vehicle main engines and the reaction control system (RCS) thrusters. Three main engines are the Shuttle OMS engines, selected for their reliability, reusability, and for storable propellants. The 24 RCS thrusters are Boeing Rocketdyne’s RS-21 engine, also chosen for the same reasons. Important parameters are listed in Table 1-2.

Table 1-2: Important ORION vehicle parameters.

OSP System / Main SP / RCS
Designation / Shuttle OMS / RS-21
mengine / [kg] / 118 / 8.4
Thrust (vac) / [N] / 26690 / 1335
Isp (vac) / [s] / 316 / 294

Both these burn Nitrogen Tetroxide and Mono-Methyl Hydrazine as their propellants and are pressurized systems (1724 kPA), fed by compressed helium. Because of these facts, their feed system and tanks are combined to save room and complexity. With the required ΔV, and ORION mass, the propellant and tank masses were calculated. All of the tanks are manufactured from a graphite composite material to maintain light weight and high strength for their structures. Significant mass values are listed in Table 1-3.

Table 1-3: Tank and propellant size and mass.

Design Concept / OSP 1
mSpacePlane / [kg] / 22757
mprop,fu / [kg] / 547
mprop,ox / [kg] / 886
mtank,prop / [kg] / 14.6
Dtank,prop / [m] / 0.36
Ltank,prop / [m] / 3.36
ttank,prop / [mm] / 2.54
mHe / [kg] / 8.6
mtank,He / [kg] / 7.2
Dtank,He / [m] / 0.7
ttank,He / [mm] / 5.9

2.0  TRAJECTORY/PERFORMANCE ANALYSIS

Vehicle 1

This vehicle was able to perform the OSP mission but not the MSP mission. Some requirements for the OSP were that it achieves 7.41˚ of crossrange using a banking maneuver and have less than 8 g’s on the astronauts. Figure 2-1 and 2-2 below show the ground track and g’s vs time respectively:

Figure 2-1: Ground Track

Figure 2-2: G-loading

Also, since no propulsive maneuver was implemented, the synergetic efficiency is 0, which is the best case scenario. So this vehicle design would be very good for the OSP mission.

Figure 2-3 below shows the total mass vs time for the MSP:

Figure 2-3: Mass vs time

Since the final mass (12145 kg) is below the initial empty weight (14721 kg), this vehicle design does not satisfy the MSP mission. Also, the synergetic efficiency is 1.24, showing that the aero-assist actually requires more ΔV than a pure propulsive maneuver would.

3.0  AEROTHERMODYNAMICS

The vehicle was broken down for aerodynamic studies using simplified shapes to represent the main features of the body (nose, leading edge, windward surface, upper body, flap) using Newtonian Theory. From this, an aerodynamic analysis was done to optimize the vehicle to have a maximum L/D while keeping the usable interior volume up and keeping the total mass down. While creating a vehicle that performed well aerodynamically, correlation had to be made with systems to create a vehicle with a range of center of gravity (CG) locations large enough that a common CG could be found between the two groups.

There were four components of drag calculated to generate a single drag value for the entire vehicle. These drag components are Newtonian drag, skin friction drag in the hypersonic and supersonic regions of flight, supersonic wave drag, and viscous drag. Lift was calculated using only the Newtonian method, thus was constant for all times during flight.

Once a vehicle design was optimized for aerodynamics, it was then analyzed to test the heating experienced during its return back to the ground. The heating subroutine used a flat plate body with a cylindrical leading edge as the model for the heating. Though incorrect, further design would elaborate on the leading edge heating and windward surface heating derivations that are in work. The vehicle Thermal Protection System (TPS) was customized by using the trajectory routine to simulate a return flight to earth. Thickness values at different sections along the body were calculated for a Reinforced Carbon Carbon (RCC) material.

The vehicles were designed to accomplish two missions. The first mission was to serve as an Orbital Space Plane (OSP) which didn’t require any specific aerodynamic capabilities other than to provide enough L/D over the entire flight range to allow for a slow and gradual descent back to earth. The second mission was to serve as a Military Space Plane (MSP) and provide the capability to use an aero-assist maneuver to change the inclination of the space plane. To do this, a large L/D was necessary. The aero-assist maneuver is performed between 40,000m and 50,000m, thus this is where the L/D capabilities need to maximized. For Design Concept 1, an approximate L/D of 4.5 was achieved at these altitudes. After analysis with Design Concept 1 in the MSP mission, it was discovered that a better L/D was needed. Design Concept 2 was designed to accomplish this and was successful; it achieves an L/D of approximately 6.5 in the altitude region of which the aero-assist maneuver would be performed.

4.0  STRUCTURES AND LOADS

The structural analysis of design concept one was made using ANSYS and ProE. Our vehicle shape is an elliptical half cone with rounded leading edges and a nose cone. ANSYS does not have many options for methods to manipulate shapes. It was then discovered that a ProE model could be imported into ANSYS to do a finite element analysis. In ProE an elliptical cone can be modeled that models our vehicle more correctly and also provides a more accurate structural analysis. The structural model is made to fit inside the outer mold line. In the structural model, but not in the actual vehicle, there are two floors. They are used to distribute the mass of the supplies inside. The structural analyses being done are of the loads at launch and the loads at parafoil deployment. For both the launch and the parafoil load conditions a stress analysis and eigenbuckling analysis were done.

Design concept one is entirely made out of high strength Graphite/Epoxy composite. This material was chosen because it has the highest strength to weight ratio compared to aluminum and titanium. It also has the highest Young’s Modulus, 5.52e10 Pa. This high of a Young’s Modulus helps to resist buckling during launch and parafoil deployment. Since this material has a very low density the mass of design concept one is lower than if it were made from aluminum or titanium.

Design concept one withstood both launch and parafoil conditions for the Orbital Space Plane mission and the Military Space Plane mission using a safety factor of 1.7.

5.0  SYSTEMS

Figure 5 - 1

ORION is 22,757 kg with an empty weight of 21,250 kg. It is the summation of all the components depicted in Figure 5-1—including, but not labeled, TPS (base, blanket) and structure. The distribution of mass is shown below as Figure 5-2.

Figure 5 - 2

It has a maximum payload capacity of 6031 kg. With accommodations for seven astronauts, the crew compartment contains 58.84 m3 of usable volume. The vehicle was designed with many considerations in mind—providing the best possible aerodynamic performance while managing the physical space required to contain a crew, fuel, payload, etc. As shown in Figure 5-3, ORION has a total length of 22.5 m, rear width of 7.2 m, and rear height of 2.791 m.

Figure 5 - 3

ORION’s center of gravity is optimally located as shown below in Figure 5-4. The propellant, OX, and Helium is oriented about the CG such that no significant change will occur as a result of fuel burn. In other words, the CG remains in this location regardless of mission time, thereby always achieving the best trimmed L/D.

Figure 5 - 4