4.1.1 Vehicle Overview 10
4.1 200g Payload
4.1.1 Vehicle Overview
The launch vehicle carrying the 200 g payload (Fig. 4.1.1.1) hitches a ride on a balloon up to an altitude of 30 km where the first of three stages is ignited. At 30 km, the rocket launches in a vertical orientation from a gondola that is attached to the balloon. Once the rocket finishes burning the propellant in all three stages, the designed orbit perigee is 486 km. When random uncertainties in vehicle performance characteristics are included in the design (Monte Carlo analysis), the launch vehicle achieves an average perigee of 437 km.
Fig. 4.1.1.1: Launch vehicle stack up – 200g payload.
(Daniel Chua)
4.1.1.1 Launch System Breakdown
4.1.1.1.1 Gondola and Balloon Components
Providing support to the launch vehicle and guidance at take-off, the gondola is an all aluminum structure. To support the launch vehicle there are three equally spaced, horizontally oriented, rings that attach to the launch vehicle’s outer structure (Fig. 4.1.1.1.1.1). Also positioned horizontally are a square frame (at the bottom of the gondola) and flange (at the top of the gondola). Connecting these rings and frame are four equally spaced, vertically oriented launch rails that guide the launch vehicle off the gondola at ignition.
Fig. 4.1.1.1.1.1: Launch vehicle and gondola configuration – 200g payload.
(CJ Hiu)
The gondola is connected to a spherical balloon, filled with helium, made of polyethylene plastic. During flight, the gondola carrying the launch vehicle is suspended below the balloon. We assume that the balloon pops right before the launch vehicle passes through it. As the balloon rises, the gas expands and the balloon is sized to hold the gas at an altitude of up to 30 km. The battery, that powers the communications with the range safety officer on the ground, is attached to the flanges of the gondola. Neither the balloon or gondola are reused. Fig. 4.1.1.1.1.2 puts the size of these components with respect to the launch vehicle into perspective.
Fig. 4.1.1.1.1.2: Size comparison of the gondola, launch vehicle, and balloon – 200g payload.
(CJ Hiu)
4.1.1.1.2 First Stage
Fig. 4.1.1.1.2.1 is an exploded view of the launch vehicle. A reference table, summarizing the sizing and propulsion information for each stage, is also provided. Please refer back to it while reading the descriptions of each stage.
Fig. 4.1.1.1.2.1: Exploded view of launch vehicle stack up and parameter summary – 200g payload.
(Stephen Bluestone, Amanda Briden, Nicole Bryan, CJ Hiu, Molly Kane, William Ling ,Sarah Shoemaker)
A hybrid first stage with a hydroxy-terminated polybutadiene (HTPB) solid fuel and hydrogen peroxide (H2O2) liquid oxidizer pairing is pressurized with gaseous nitrogen and provides a thrust of 34 kN. Part of the first stage propellant is tapped off to support the liquid injection thrust vector control (LITVC), which is used to steer the rocket. Made out of light-weight space-grade aluminum, the structure can withstand a maximum acceleration of 4.54 Gs. The first stage is 70.11% of the launch vehicle’s gross liftoff mass (GLOM) and the length of this stage is 6.86 m. Fig.4.1.1.1.2.2 is a dimensional drawing of the first stage.
Fig. 4.1.1.1.2.2: Dimensional drawing of the first stage – 200g payload.
(Nicole Wilcox)
4.1.1.1.3 Second Stage
The second stage is an ammonium perchlorate (AP), aluminum (Al), and HTPB solid motor with an extra tank of H2O2 to provide fuel for the LITVC. The H2O2 is again pressurized with gaseous nitrogen. This stage imparts a thrust of 8.8 kN. Able to withstand a maximum acceleration of 4.87 Gs, the second stage is made of space-grade aluminum. A cone truncated in the mid-section is used to connect one stage diameter to the next such that there are no gaps in the structure; this is called a skirt. The most significant part of the avionics package is located on the interior of the skirt connecting the second and third stages. The avionics package located in the skirt includes a battery, telecom, central processing unit (CPU), and CPU equipment. These features increase the avionics mass from the first by a factor of 5, for a total avionics mass on the second stage of 30 kg. The second stage is 27.87% of the launch vehicle’s GLOM. This stage is
3.11 m long and a dimensional drawing is shown in Fig. 4.1.1.1.3.1.
Fig. 4.1.1.1.3.1: Dimensional drawing of the second stage – 200g payload.
(Nicole Wilcox)
4.1.1.1.4 Third Stage
Since the avionics is jettisoned along with the second stage at the end of its burn, we spin the third stage of the launch vehicle to maintain stability. The propellant type and structural material are identical to the second stage. The third stage is 2.02% of the launch vehicle’s GLOM. Stage three is 1.06 m in length and a dimensional drawing follows in Fig. 4.1.1.1.4.1.
Fig. 4.1.1.1.4.1: Dimensional drawing of the third stage – 200g payload.
(Nicole Wilcox)
4.1.1.1.5 Nose Cone Component
The nose cone protecting the top of the launch vehicle from extreme heating is made of aluminum and titanium. An additional feature of the nose cone is a blunted tip made of titanium, which is a heat resistant material. The nose cone is jettisoned once the vehicle reaches an altitude of 90 km (out of the Earth’s atmosphere). The nose cone jettison occurs prior to the separation of the first stage.
4.1.1.2 Mission Requirements Verification
What are the chances that we reach an orbit with a periapsis of at least 300 km?
There is a 99.99% chance that our launch vehicle reaches a periapsis of 300km. After 10,218 Monte Carlo simulations launch vehicle only fails once (Fig. 4.1.1.2.1). We therefore meet the mission requirement of 99.86% success rate, considering only non-catastrophic failures. An average perigee, shown as the peak of the histogram in Fig. 4.1.1.2.1, of 437 km is achieved.
Fig. 4.1.1.2.1: 200g periapsis altitude histogram with std = 21.803 km and mean = 437.44 km.
(Alfred Lynam)
What are the chances of a failure that results in complete loss of mission?
Accurately predicting the mission success rate, including failures that result in complete loss of mission, is difficult to do without built and tested hardware. Therefore, we turn to the historical success rates of the Ariane IV, Ariane V, and Pegasus, to predict ours. We use the success pattern of Pegasus as it is the only vehicle is air-launched. We predict a 93.84% success rate, which includes catastrophic failures and is achieved after 24 launches.
4.1.1.3 Mission Timeline - A Launch in the Life of the 200g Payload Launch Vehicle
T- 1:35:42 to launch
The entire launch system begins its 1 hour and 35 minute ascent to its launch altitude of 30 km. On average, the system drifts 121 km before reaching the launch altitude. Prior to ignition, a range safety officer on the ground checks the status of the launch system and has the authority to proceed with or abort the launch. Fig. 4.1.1.3.1 is a visual representation of the stages of flight described in the timeline.
T+ 00:00:00 to launch – We are go for launch!
If all systems are go, the first stage is ignited and the launch vehicle is guided off the gondola via four launch rails. We assume that the balloon pops as the launch vehicle passes through it. Throughout the course of the burn, the position of the launch vehicle is determined at every instant by the control system which follows a near optimal steering law. During the first stage the launch vehicle climbs out of the atmosphere and jettisons the nose cone.
T+ 00:02:17 First Stage Burn-out
Approximately two thirds of the way through the first stage burn, the launch vehicle begins a pitch over maneuver. This initial maneuver is of the same form of that used in the Apollo program. After burning for 136.8 s and climbing to an altitude of 97.4 km, the first stage separates.
T+ 00:05:35 Second Stage Burn-out
During this phase, the launch vehicle continues to pitch over to burn off velocity in the radial direction. At orbit insertion, radial velocity needs to be zero in order for a circular orbit to be achieved. With the burn duration of 207.7 s and burn out altitude of 347 km, the second stage separates, jettisoning the bulk of the system’s avionics.
T+ 00:08:56 Third Stage Burn-out – We’re in orbit!
For the duration of the third stage burn, the launch vehicle uses spin stabilization to maintain its orientation and does not require avionics control or LITVC. This means that the vehicle’s orientation from the end of the second stage burn through the third is maintained. After a 191.9 s third stage burn time, the launch vehicle ends its ascent and enters an orbit with a perigee of 437 km. The total mission time is 1.7 hours.
Fig. 4.1.1.3.1: Mission Timeline – 200g payload.
(Amanda Briden, Kyle Donahue, Jeffrey Stuart)
Author: Amanda Briden