conceptual design document (cdd)

University of ColoradoDepartment of Aerospace Engineering Sciences

ASEN 4018

Conceptual Design Document (CDD)High Altitude Lifting Orbiter

(HALO)

Monday 30th September, 2019

1. Information

1.1. Project Customers

Name: Dale LawrenceEmail: [email protected]: 303-492-3025

1.2. Team Members

Name: Braden BarkemeyerEmail: [email protected]: (970) 729-0545Position: Altitude and Orbit Lead

Name: Nolan FergusonEmail: [email protected]: (720) 289-5349Position: Systems Engineer

Name: David CeaseEmail: [email protected]: (203) 980-0832Position: Project Manager

Name: Ryan LansdonEmail: [email protected]: (303) 802-7260Position: Electronics Lead

Name: Kelly CrombieEmail: [email protected]: (630) 577-7412Position: Software Lead

Name: Jacob MarvinEmail: [email protected]: (303) 995-7565Position: Safety Lead

Name: Jared DempewolfEmail: [email protected]: (303) 718-7913Position: Testing and Verification

Name: Kyle McGueEmail: [email protected]: (843) 697-8747Position: Environmental Lead

Name: Tyler FaragallahEmail: [email protected]: (720) 315-6513Position: Manufacturing Lead

Name: Paolo WilczakEmail: [email protected]: (303) 960-2756Position: Chief Financial Officer

Contents

1 Information 11.1 Project Customers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.2 Team Members . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

2 Project Description 42.1 Project Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42.2 Objectives . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52.3 Concept of Operations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62.4 Functional Block Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 72.5 Functional Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

3 Design Requirements 9

4 Key Design Options Considered 124.1 Lifting Orbiter Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

4.1.1 Wing with Mounted Propeller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 124.1.2 Lifting Wing on a Rotational Tether . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 144.1.3 Separate Orbiter and Altitude Control Motors . . . . . . . . . . . . . . . . . . . . . . . . . . 15

4.2 Thermal Control Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 164.2.1 Passive Thermal Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 174.2.2 Active and Passive Thermal Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

4.3 Tether Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 184.3.1 Static Fixed-Length Tether . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 184.3.2 Passively Unspooled Tether . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 194.3.3 Dynamically Unspooled Tether . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

4.4 Venting Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 214.4.1 Solenoid Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 224.4.2 Electric Motor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

4.5 Balloon Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 244.5.1 HAB 1000 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 264.5.2 HAB-TX-1500 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 264.5.3 HAB 3000 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

5 Trade Study Process and Results 275.1 Lifting Orbiter Design Trade Study . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

5.1.1 Metrics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 275.1.2 Scoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28

5.2 Thermal Control Design Trade Study . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 295.2.1 Metrics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 295.2.2 Scoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30

5.3 Tether Design Trade Study . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 305.3.1 Metrics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 305.3.2 Scoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32

5.4 Venting Control Trade Study . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 335.4.1 Metrics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 335.4.2 Scoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34

5.5 Balloon Selection Trade Study . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 345.5.1 Metrics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 345.5.2 Scoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35

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6 Selection of Baseline Design 366.1 Lifting Orbiter Design Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 366.2 Thermal Options Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 366.3 Tether Design Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 366.4 Venting Options Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 366.5 Balloon Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 366.6 Baseline Design Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36

Nomenclatureα Angle of attack [rad]

ω Rotational speed [ rads ]

ρ Density [ kgm3 ]

A Area [m2]

CD Coefficient of drag

CL Coefficient of lift

D Drag force [N]

E Energy [J]

F Force [N]

f Frequency [s−1]

k Thermal Conductivity [ WmK ]

L Lifting force [N]

m Mass [kg]

P Power [W]

Q Heat [J]

r Radius [m]

S Airfoil area [m2]

T Temperature [K]

t Time [s]

V Velocity [ ms ]

Vb Battery Voltage [V]

A

Volume [m3]

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2. Project DescriptionMission Statement: The High Altitude Lifting Orbiter (HALO) mission seeks to innovate the process of collecting

turbulence data in the stratosphere. This will be accomplished by designing a tethered payload system which will beable to autonomously execute controlled orbits around the wake of a high altitude balloon. Additionally, this orbitersystem will be actuating controlled altitude oscillations between 25 km and 35 km while recording and transmittingtemperature, pressure, IMU, and GPS sensory data in order to develop a greater understanding of turbulence in thetarget region.

2.1. Project Overview

Currently, there is limited knowledge of turbulence patterns present in the stratosphere. This is largely due tothe fact that current turbulence measurement methods are inefficient and expensive. Recently, the Air Force Officeof Scientific Research (AFOSR) funded the Hypersonic Flight in the Turbulent Stratosphere (HYFLITS) [13] projectto collect turbulence data in this region of the atmosphere. Collection of this data will allow for more informeddevelopment of high altitude aircraft that operate in this region. HYFLITS utilizes a single-use balloon system tocapture turbulence data between altitudes of 20 km and 35 km for roughly 1.5 hours. This data collection methodis limited by the balloon’s wake; turbulence data cannot be collected during ascent because the turbulence sensoris located directly within the wake of the balloon. The High Altitude Lifting Orbiter (HALO) project will improveupon this mission by moving the sensors out of the wake and implementing multiple controlled ascents and descentsbetween 25 km and 35 km. These improvements will allow for increases in turbulence data collection per balloonlaunch. Increasing the quantity of turbulence data collected per balloon launch is significant because the balloon andsensor package are designed for single time use only. This increase in turbulence data collection will be desirable solong as the useful-data per dollar ratio is increased.

Figure 1a displays the flight profile of the HYFLITS mission, while Figure 1b displays the proposed flight profileof the HALO mission. Figure 1b shows three ascent and descent data collection phases, which would produce sixtimes the amount of data produced by HYFLITS. This is a tentative number of data collection phases that may bereevaluated throughout the design process.

Figure 1. Proposed Flight Profile Changes

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2.2. Objectives

Success for project HALO will ultimately rely on achieving two main goals: developing a tethered orbiter systemto remove the measurement devices from the balloon’s wake, and controlling lift and drag on the tethered orbiterin order to oscillate balloon altitude within a target range of 20 km - 40 km. This orbiter is defined as the tetheredsystem which will orbit below the high-altitude balloon outside of the balloon’s wake. Accomplishing these goals willincrease turbulence data collection efficiency by means of lengthening the data collection period per flight and thenumber of ascents and descents per flight. In addition to these goals, a sufficient and safe mission termination protocolwill be initiated at the end of the mission to prevent a wayward balloon. The termination method from the HYFLITSmission may be integrated initially. However, HALO will seek to improve on this method by implementing a systemtriggered by sensory measurements. Additionally there will need to be a fail-safe method to arrest the balloon andpayload in case of a balloon burst. With this in mind, Table 1 was constructed, which elaborates upon the requirementsabove. In the following paragraphs, each level of success will be outlined in detail.

Level I requirements represent basic concepts that will need to be accomplished and tested in order to justify abase level of success. At this level, the tethered orbiter will be able to orbit beneath the balloon while ascending to abaseline altitude of 20 km. During this time, Inertial Measurement Unit (IMU) data will be recorded and transmittedto the receiving antenna on top of the CU Aerospace building so as to confirm a successful payload orbit around theballoon. Thermal systems will be implemented in order to ensure that all electronics as well as the power supplyremain within their operational temperatures throughout flight. Temperature data will also be transmitted in order toverify the success of these thermal systems. The balloon will be neutrally buoyant at this baseline altitude; this willsuggest that a change in lift or drag will easily change balloon altitude in preparation for level II. The balloon andpayload will be capable of controlled descent after reaching the target altitude. Additionally, there will be a passivearresting device to mitigate hazards due to unanticipated balloon burst.

Level II will then implement controls such that the tethered orbiter can bring the balloon to a maximum speed of2 m

s in ascent or descent. The balloon will reach its baseline altitude of 20 km, descend to 15 km, and ascend up to25 km. All altitudes will be allowed an uncertainty of 1 km. At this level IMU, turbulence, temperature, pressure, andGPS data will be collected during the initial ascent, through the oscillation, and during the final descent. This GPSdata is what will allow for reaching targeted altitudes during controlled ascents and descents. Controlled descent willnow be triggered autonomously, based on sensory measurements. At some point during flight, the tethered orbiter willdemonstrate the capability to remove itself from the balloon wake. This balloon wake will initially be defined by theshadow area of the balloon, and may change due to high wind velocities or further investigation.

At Level III the tethered orbiter will be fully out of the balloon wake. The balloon will reach a baseline altitudeof 30 km and immediately begin oscillating between 25 km and 35 km. IMU, turbulence, temperature, pressure, andGPS data will be collected during the initial ascent, throughout each oscillation, and during the descent. These targetaltitudes were chosen per the customer’s request, and may be subject to change throughout the design stage. As inlevel II, end-of-life termination will be triggered autonomously based on sensory measurements.

Each of these design aspects will be sufficiently tested on the ground prior to launch. The tethered orbiter willbe tested by measuring its potential orbit radius on the ground and comparing it to predicted balloon radii at thevarious altitudes of interest. Its altitude control will be tested by demonstrating that it can produce sufficient lift anddrag in near-flight conditions. Data collection will be tested by transmitting and receiving data at a significant distance.Termination will be tested and confirmed as triggered by the desired stimulus, under near-flight conditions. Electronicsand thermal systems will be tested by maintaining operational temperatures in simulated near-slight conditions.

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Tethered Orbiter Altitude Data Collection Descent ArrestLevel I The payload shall or-

bit below the balloon.This orbit will be stableand periodic.

The balloon and pay-load will achieve a tar-get altitude of 20 kmand be neutrally buoy-ant. The balloon willremain at this altitudeuntil termination.

The orbiting payloadwill contain an IMU toverify a successful or-biter trajectory. It willalso contain a tempera-ture sensor to verify asuccessful thermal sys-tem. This data will betransmitted to a receiv-ing antenna through-out the flight duration.All sensors will beoperational throughoutflight.

System will include afail-safe descent arrest-ing device in the eventof balloon burst. Afterreaching the target al-titude, the balloon willenter a controlled de-scent at a maximumspeed of 5 m

s .

Level II The payload will be ina stable orbit below theballoon, demonstratingthe ability to removeitself from the wakeat some point duringflight. It will alsogenerate enough lift ordrag to bring the bal-loon and payload toa velocity of 2 m

s ±

20% in ascent and de-scent. Orbiter speedwill not introduce tur-bulence measurementbias.

Controls will be imple-mented that will allowthe tethered orbiter tochange the balloon andpayload altitude. Theballoon will reach 20km, descend to 15 km,ascend to 25 km, andterminate. GPS mea-surements will be usedto approximate the al-titude of the balloon.These altitudes will beallowed an uncertaintyof 1 km.

Turbulence, tempera-ture, GPS, IMU, andpressure data willbe collected on thepayload throughoutits flight. This datawill be transmitted toa receiving antennaduring flight. Turbu-lence sensors will beoriented towards therelative wind in orderto minimize bias.

Controlled descentwill be initiatedautonomously as trig-gered by a sensorymeasurement.

Level III The stable payloadorbit will remain fullyoutside of the balloonwake for the entireflight, ignoring occa-sional perturbationsdue to wind gusting. Italso must be capable ofgenerating lift or dragto bring the balloonsystem to a velocityof 2 m

s ± 20% duringascent or descent.

The payload will reacha target altitude of 30km and then oscillatebetween 25 km to 35km until power de-pletion. GPS mea-surements will approx-imate the altitude of theballoon. These alti-tudes will be allowedan uncertainty of 1 km.

Table 1. Levels of Success for the HALO

2.3. Concept of Operations

The following figure (Fig. 2) shows the Concept of Operations for the final launch day of the HALO balloonsystem. Weeks before launch, the HALO team will coordinate with administrators in charge of the CU Aerospace

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building receiving antenna in order to find a suitable launch time. Prior to launch, the HALO team will first observewind vectors and evaluate launch safety; if there are high winds blowing towards Denver International Airport orother high-traffic zones, launch will be rescheduled. Figure 2 displays the launch procedure after this safety check iscomplete. System setup will require at most two people. The balloon will be filled with the desired amount of heliumafter the tether is attached, and must be released by a single person. While this launch is not technically regulatedby the FAA due to the small size of our balloon, it is required to meet FAA guidelines per customer request. Thetether will extend after liftoff for ease of launch. Once achieving neutral buoyancy, the balloon system will go throughautonomous ascent/descent cycles for data collection. In order for the launch to be successful, the system must be ableto communicate with the ground system by transmitting live sensory data. The flight will then be terminated, triggeredautonomously by an on-board sensory measurement.

Figure 2. Concept of Operations for HALO Mission

2.4. Functional Block Diagram

Contained below in Figure 3 is the Functional Block Diagram (FBD) for the HALO mission. Three systems for theHALO mission are to be developed or expanded upon. This includes the orbiting system, the altitude control system,and the termination system. In its current state, the termination system consists of a venting servo that is activatedupon reaching a designated target altitude after a single flight cycle. HALO is proposing a new termination systemthat is dependent on sensory information and will include additional safety measures. The orbiting system consistsof control software and a propulsion method. The altitude control system consists potentially of a lifting body, actingas the main body for the orbiting payload. As a target altitude is reached, indicated by GPS data, the altitude controlsystem conducts a controlled descent/ascent from its current position to the desired altitude.

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Figure 3. Functional Block Diagram for HALO Mission

2.5. Functional Requirements

FR 1.0: Balloon and its payload must meet all FAA [6] guidelines, including those subjecting the balloon and itspayload to designation as an unregulated balloon.FR 2.0: Launch setup must be able to be conducted by a maximum of two people, and launch must be able to beconducted by a single individual.FR 3.0: There will be a method of controlling the altitude of the balloon and payload.FR 4.0: The payload must be capable of orbiting beneath the balloon.FR 5.0: The payload will be able to transmit collected data to the ground receiver.FR 6.0: The balloon will terminate flight resulting in a controlled descent upon mission completion.FR 7.0: As specified by the customer, the balloon must contain a controlled descent arrest system (CDAS) to preventharm to property or person upon balloon burst.FR 8.0: Project HALO will cost no more than $5,000 to develop, and no more than $2,000 per flight when componentsare purchased in quantities of 50.

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3. Design RequirementsFR 1.0: Balloon and its payload must meet all FAA guidelines, including those subjecting the balloon and itspayload to designation as an unregulated balloon. In addition, Balloon and payload shall meet specificationsset forth by customer.

Motivation: As requested by the customer, in order to keep the balloon and its payload launchable by smallteams and without significant regulationafrom the FAA, unmanned free balloon regulations (FAA Part 101) must beconsidered as the ultimate design specifications.

Verification: The following design requirements will be met. Re-verification of these policies with those familiarwith FAA balloon guidelines will be undertaken.

• DR 1.1: The payload will be weight constrained.Motivation: Any payload above the weight classes designated below subjects the balloon to a title of an Un-manned free balloon, which requires the payload and balloon to be further regulateda.

Verification: Weights of components for each system will be tracked. Final weight testing will be undertakento ensure payload weight restriction met (if DR 1.1.2 is met, surface area will be increased to reduce the smallestsurfaces of the payload package).

– DR 1.1.1: The payload will be under four pounds, OR

– DR 1.1.2: If the payload consists of a single package, and weighs between 4 to 6 pounds, the payloadmust have a weight to size ratio of less than three ounces per square inch on any surface of the package(determined by dividing the total weight in ounces of the payload package by the area in square inches ofits smallest surface), OR

– DR 1.1.3: If the payload consists of two or more packages, these must weigh under six pounds individu-ally.

• DR 1.2: Tether or other suspension device must NOT be able to withstand an impact force of 50 lb.Motivation: Any payload with tether or other suspension device with a impact strength greater than 50 lb isclassified as a Unmanned free balloon, which subjects the payload to further regulationa.Verification: Strength of tether or suspension device will be tested to ensure the ability to withstand weight ofpayload, and break with an impact force of 50 lbs.

• DR 1.3: Lifting gas of balloon will be Helium.Motivation: As specified by the customer, the lifting gas will be helium.Verification: Balloon will not be subject to testing or filling of other lifting gases, primarily Hydrogen.

• DR 1.4: Payload shall not be toxic to environment.Motivation: The payload, as specified by customer, will not be toxic or harmful to the environment, in theevent of no payload recovery.Verification: No toxic materials (such as lead weights) will be added to payload or balloon.

FR 2.0: Launch setup must be able to be conducted by a maximum of two people, and launch must be able tobe conducted by a single individual.

Motivation: As required by the customer, in order for the developed payload and balloon system to remainmodular and transportable by small teams in varied terrains and locations, it is necessary that the system does notprovide unreasonable effort to setup and launch.

Verification: Dry tests and launches will be conducted and timed by 2 or less individuals before a final flight test.

• DR 2.1: Weight of payload must be able to be handled by one individual.

Motivation: For launch to be physically conducted by a single individual, the payload must be of reasonableweight. This allows the individual to handle both the balloon and the payload for a given launch.

Verification: Dry tests and launches will be conducted and timed to ensure that a single individual can managea launch.

aRegulation, as it is used here, pertains directly to further design specifications or safety features that must be implemented according to FAAlaw.

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• DR 2.2: Tether or suspension devices must not be longer than 10 meters at launch.

Motivation: In order for launch to be conducted smoothly, with a reasonable and non-destructive accelerationof the payload, the length of the tether at launch must be under 10 meters, as specified by the customer.

Verification: Tether will remain under 10 meters. If devices are used to extend the tether length in flight, thenthese devices will be tested in stationary ground tests, a moored test, and finally a flight test.

FR 3.0: There will be a method of controlling the altitude of the balloon and payload.

Motivation: Altitude control is necessary to accommodate data collection periods throughout the flight, whetherby a single ascent/descent, or by multiple ascents/descents.

Verification: The payload will contain a device to autonomously control altitude, as well as a GPS to confirmposition of collected data.

• DR 3.1: Balloon system altitude ascent and descent speed maximum speed shall be within 20% of 2 m/s.

Motivation: As requested by the customer, speed must remain relatively low (within 20% of 2 m/s) as not todisturb data collection received by instruments.

Verification: Moored testing will be undertaken to ensure that in a data collection phase, speeds are within anacceptable range. In addition, some in flight method of velocity verification will be recorded.

• DR 3.2: Balloon system must be able to achieve neutral buoyancy to some capacity at a specified altitude.

Motivation: In order to improve the current data collection capacity, it is necessary that neutral buoyancy isreached. Neutral buoyancy will allow the designed method of altitude control to be effective for cycles of datacollection. It will allow the orbiter to increase its lift and change the altitude of the balloon.

Verification: Modeling of a similar system will be conducted on the ground. On board devices will confirm theacceleration or speed of the balloon/payload system, and corrections will be made to achieve neutral buoyancy.

FR 4.0: The payload must be capable of orbiting beneath the balloon.

Motivation: As a proof of concept, it is necessary that orbiting beneath the balloon must be undertaken for theability to collect data outside the wake of the balloon. This will allow extended data collection periods throughout alarge majority of the stratosphere.

Verification: Static tethered testing around a stationary pole will be conducted before a moored test, followedby a flight test. This will allow testing of an orbital system on the ground, a controlled environment in the air, and asimulated mission.

• DR 4.1: Payload must have some ability to determine the position of the orbit and position of orbit relative tothe center of the balloon.

Motivation: In order to determine the stability of the orbit, as well as to aid in determination of whether theorbiting payload is collecting adequate data, it is necessary to determine the position in the orbit.

Verification: Static tethered testing around a stationary pole will be conducted before a moored test, followedby a flight test.

• DR 4.2: Orbiting system or device must not damage or deform balloon.

Motivation: Data collection must not interfere with the balloon flight, in order for the mission to proceedsafely and effectively.

Verification: The mechanisms and devices conducting the orbiting will be tested on stationary tether lines.These mechanisms must not cause harm to the balloon.

FR 5.0: The payload will be able to transmit collected data to the ground receiver.

Motivation: While the primary goal is to institute an orbiting payload for data collection outside the balloonwake, this data will need to be transmitted and characterized for turbulence modeling.

Verification: A final flight test will be undertaken to verify the successful collection of data from the groundstation located on the Aerospace building.

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• DR 5.1: Data transmission must be collectible within a 100 km range of the stationary ground station.

Motivation: The range was designated to be 100 km by the customer to ensure the transmission of the datawhile taking into account balloon altitude gain and lateral drift from the ground station.Verification: Varied altitude testing will be undertaken, utilizing the varied terrain of Boulder and the sur-rounding area to test transmission capabilities at different altitudes. A moored ground test will be conductedprior to a flight test as well.

FR 6.0: The balloon will terminate flight resulting in a controlled descent upon mission completion.

Motivation: As required by the customer, the system will self execute into descent so that the balloon is not freefloating in the air space.

Verification: Some method of verification will be transmitted to confirm the success of the termination system inflightb This system will consist of two methods of termination to some degree, for redundancy.

• DR 6.1: If payload is to make soft landing in water, the payload shall self-scuttle.Motivation: Removes the possibility of the payload becoming floating debris for waterborne vehicles.

Verification: Testing of payloads sinking capabilities will be conducted.

FR 7.0: As specified by the customer, the balloon must contain a controlled descent arrest system (CDAS) toprevent harm to property or person upon balloon burst.

Motivation: As specified by the customer, the balloon must contain a controlled descent arrest system (CDAS)to prevent harm to property or person upon collision with earth after balloon burst.

Verification: Speed and force impact testing will be conducted to ensure a safe descent speed, and a safe impactmomentum. This will be verified via a drop test off of the aerospace building.

• DR 7.1: The CDAS will passively deploy upon balloon failure or flight termination.Motivation: The CDAS must be activated to establish a safe descent speed regardless if activation is passive orintentionally triggered.

Verification: Drop testing will be carried out to ensure that CDAS activates passively. Ground based testingand controlled flight testing (Moored) will be undertaken to ensure that the system activates successfully priorto flight testing.

• DR 7.2: The CDAS must provide a descent speed of a maximum of 5 m/s c .Motivation: After termination, the payload and balloon must descend at a safe speed, as not to cause harm tobody or property.

Verification: Accelerometer testing with the CDAS system will be undertaken prior to flight testing to verifydescent speed. Ground based testing and controlled flight testing (Moored) will be undertaken to ensure that thesystem activates successfully prior to flight testing. Force impact testing will also be conducted to ensure thatminimal damage is caused by any falling components.

FR 8.0: Project HALO will cost no more than $5,000 to develop, and no more than $2,000 per flight whencomponents are purchased in quantities of 50.

Motivation: As required by the customer, these cost values are set to allow multiple launches without needing amassive budget.

Verification: Costs of each component will be tracked and summed to ensure that the project stay low in cost.

bThis will be done if within transmission range after the designated flight time.cThe value of 5 m/s was provided by the customer as a estimate. However, more research of a safe descent speed will be conducted for the

eventual payload weight.

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4. Key Design Options Considered4.1. Lifting Orbiter Design

The main objectives of this project are to record data outside of the balloon’s wake and to control multiple au-tonomous ascents and descents between 25 km and 35 km. In order to achieve these objectives, a system which orbitsbelow the balloon must first be developed. Additionally, there must be a method of autonomously controlling theoverall altitude of the balloon based on sensory measurements. The orbiter must be capable of providing a force greatenough to move out of the balloon wake, and the lifting capabilities must be great enough to bring the entire balloonsystem to an ascent or descent rate of roughly 2 m

s . This will prove to be difficult due to the large drag force caused bythe large surface area of the balloon at high altitudes. Further, it is critical that the sensory measurement devices aremoving in a stable manner as to corrupt any recorded turbulence data. The following sections seek to explore designscapable of achieving these feats.

4.1.1. Wing with Mounted Propeller

With this design, the orbiter will take the shape of an airfoil with a mounted propeller. In conjunction with a tetherattachment, this will allow the airfoil to maintain circular motion around the wake of the balloon. Further, the angleof attack of this airfoil will be autonomously controlled in order to affect lift forces on the entire balloon system.This autonomous control will be actuated by a servo that will change the airfoil angle of attack based on altitudemeasurements from the GPS device. In order to ascend, the airfoil will increase the angle of attack from level to adesired angle to increase the lift on the payload. After the peak is reached, the wing would rotate to a negative angleof attack and provide negative lift, taking the balloon down to the valley of the flight profile.

The lift generated from this airfoil, as well as airfoils described in later designs, will be approximated usingEquation 1.

L =12ρV2S CL (1)

In this equation L is the lift generated by the airfoil, CL is the lift coefficient of the wing, ρ is the ambient air density,V is the airfoil speed, and S is the airfoil area. Since air density at altitude may be less than 1% of air density on theground [5], it will be difficult to generate this lift. Similarly, the drag on the balloon will be needed to be accounted forwhich can be found through Equation 2 below.

D =12ρV2S CD (2)

To find the velocity that the airfoil will need to fly at to maintain a circular motion, Equation 3 can be used inconjunction with the horizontal force component of the tether to balance the forces. The propeller will provide theneeded thrust to balance the wing drag and prevent orbit decay.

Fradial = mpayloadV2

rwake(3)

In this design and the final design, propellers are being used. In order to conduct a proper force analysis, the thrustof the propeller must be analyzed. Equation 4 will be used to solve this problem where r is the radius of the circlemade by the blades, V∞ is the free-stream velocity seen by the front of the wing, and ∆V is the change in velocity ofthe air from one side of the propeller to the other.

Fprop = πr2ρ(V∞ +

∆V2

)∆V (4)

A series of calculations were conducted in order to determine the required area for the wing, and it was found thata large wing area would be required for this design. In order to ascend or descend at 2 m

s , assuming a tangential orbitalvelocity of 10 m

s , a wing area of approximately 0.69 m2 will be required. Detailed calculations are located in section5.1 and further in Appendix A.

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Figure 4. Wing with Mounted Propeller Design

Figure 5. Wing with Mounted Propeller FBD

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Pros ConsUses less power for lift Difficult to initially acquire orbit

Can be modelled similarly to a plane Will require a complex control systemAll components are located in one place Hard to generate lift / requires large wings

Table 2. Wing with Mounted Propeller Pros and Cons

4.1.2. Lifting Wing on a Rotational Tether

This design is centered around a high mounted motor. A tether will extend from the motor to the airfoil, actu-ating the necessary rotational motion of the orbiter. Similar to the previous design, the angle of attack will then beautonomously controlled by a servo in order to generate positive or negative lift with regard to altitude measurements.The main advantage of this design is that wind perturbations will have a less significant effect; the spinning motor willrecover circular motion much more quickly than the wing with a mounted propeller.

The main difficulty with this design is that the high mounted motor will naturally twist the balloon throat; it willbe necessary to determine that this twisting will not significantly harm the balloon.

Similar to the first design, Equations 1-3 will be utilized when approximating the necessary orbiter rotationalspeed.

Figure 6. Lifting Wing on a Rotational Tether Design

Pros ConsEasy to restart motion if perturbed by wind Complex control system

Easily modelled Requires large wingsNo propeller analysis Motor will be high mass

May twist balloon throat

Table 3. Lifting Wing on a Rotational Tether Pros and Cons

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4.1.3. Separate Orbiter and Altitude Control Motors

This design is a direct continuation of the previous design. The primary difference is that this design implementsa separate lifting force via a tethered propeller and motor system. This propeller and motor system will be tethered tothe high mounted motor as shown in Figure 7 which will eliminate the need for lift force from the orbiter. This verticaltether could be replaced by a solid rod if deemed appropriate. The second tether will attach the orbiting payload tothe high mounted motor as in the previous design. The motor will rotate the tether which will then move the payloadoutside the wake in a circular orbit.

This design comes with significant trade offs. Its biggest advantage is that the lifting source is acting directlyupwards/downwards for more lift. The airfoil in this design is primarily for stability, and is not required for liftgeneration. This, in turn, means that the wing will not need to be nearly as large as in the previous two designs.This propeller and motor system will unfortunately contribute significantly to the mass budget, and must be evaluatedthoroughly for feasibility. If necessary, mass could be reduced by removing the airfoil and only orbiting the sensorsuite.

Similar to the first design, Equation 4 will be used in order to approximate the force generated by the proposedpropeller. Figure 7 displays the proposed Separate Orbiter and Altitude Control System design. A wing is displayed inthis figure, and is not designed to generate lift. If deemed necessary, however, it could generate lift by autonomouslycontrolling its angle of attack, as in the previous designs.

Figure 7. Separate Orbiter and Altitude Control System Design

Pros ConsSeparate components can be simpler to work with Two motors could yield a high mass

Primary lifting force parallel to Z axis High power consumptionEasy to start orbiting and rising Tethers could intersect and tangle

Table 4. Separate Orbiter and Altitude Control System Pros and Cons

In conjunction with this design, preliminary research on different motor types was conducted: specifically, high

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kV and low kV motors. Their brief specification for further investigation are located below.

High kV Motor Low kV MotorFaster RPM w/small props Low RPM w/larger prop

Low Torque High TorqueMass Heavy Light

Table 5. Motor Choice

4.2. Thermal Control Design

Thermal control is critical to the function of the payload and orbiter system. The mission profile includes travelingto altitudes where temperatures drop as low as -56.50°C [5]. The most pressing requirement of the thermal system is tomaintain a high enough temperature for any batteries to remain functional. Batteries fail at low temperatures becausethey rely upon chemical reactions that begin to react too slowly to generate a functional current at extremely coldtemperatures. The temperature that this malfunction occurs at is different for different batteries. The second drivingfactor for thermal control is that some sensors only produce reliable data within a limited temperature range. Thefollowing designs will seek to ensure that all electronics remain entirely operational throughout the data collectionphase of flight.

Figure 8. Current Payload Design With Passive Foam Insulation

Qloss =2πL(Ti − T0)ln( r0

ri)/k f oam

=2πL(Ti − T0)

ln( thickness+riri

)/k f oam= Qgen (5)

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Symbol Meaning ValueQgen Heat generated by electronics within payload 3 WattsQloss Heat loss from inside payload to outside atmosphere Steady heat transfer so equal to Qgen

L Height of cylinder 20 cmTi Temperature inside payload -20°CTo Atmospheric temperature outside payload -60°Cri Inner radius of payload 5 cmr0 Outer radius of payload r0 = ri + thickness

k f oam Thermal conductivity of average foam 0.030 WmK

Thickness Thickness of foam required to achieve proper insulation Solved for: 3.2655 cm

Table 6. Assumed Values to Test for Validity of Foam Insulation

The first step in testing the viability of a thermal control system is to determine how much material/heat wouldbe needed to keep the temperature inside the payload at a working battery level. This temperature was conservativelydetermined to be -20°C [2] based on lithium battery data which will likely be the battery type used for the HALOmission. A simple heat transfer model (located in Appendix A) was created in MATLAB that takes into account thethermal conductivity of the foam materials and the theoretical heat generated by the electronics within the payload(from either preexisting electronics or a heater) to test whether foam materials could be viably used for insulation.This model assumes a steady state heat transfer where the payload is modeled as a cylinder with no heat loss fromthe top or bottom (Figure 8). This model was used to predict the thickness of foam insulation required to keep theinternal temperature at the target temperature noted above. Equation 5 and Table 6 above illustrate the equation andvariable values used in the model. The thermal conductivity of the foam was estimated to 0.030 W

mK based on theaverage thermal conductivity of the materials being considered [8]. The electronics heat dissipation was chosen to be 3W. This value is very conservative based on a NASA article that examines the thermal systems of small spacecraft [15].Next, the atmospheric (T0) and payload (Ti) temperatures were chosen based on conservative estimates discussed inthe preceding paragraph [2] [5]. The rest of the sizing values (L, ri) were chosen based on estimates of our final payloadsize. For this reason, some of the values are rather arbitrary. However, they would not be drastically different fromthe values that will actually be used in our mission, and thus remain valid for a design validity check. This modelproves that foam material can function as a proper insulator for the mission because the thickness required was foundto be 3.2655 cm with an electronic heat generation of 3 W. This is a small enough thickness that it could easily beimplemented on the orbiter without being too bulky. The thickness could also be reduced by adding a heater or if theelectronics end up producing more heat than 3 W.

It is important to note that a method of pure active thermal control was not considered because it is not feasible.Any material used to cover the payload bus will function as an insulator and thus add passive control. For this reason,only passive control and a combination of active and passive control will be considered as viable designs.

4.2.1. Passive Thermal Control

Passive thermal control is the simpler of the two designs and relies entirely on passive insulation and heat generatedfrom essential electronics that have functions separate from heating. An example of these electronics would be anelectric motor for the orbiter that drives a propeller. The only types of passive insulation considered for this particularmission have foam-like compositions. Only foam-like insulations were considered as viable options because they havelow thermal conductivities and low densities.

One pro of this design is that it would be very simple and there would be less moving parts and electronics withinthe payload. Another pro is that there would be a smaller chance of catastrophic failure with passive thermal controlas opposed to active thermal control. The heater required for active thermal control could malfunction and either heatthe payload to too high of temperature or fail to warm it enough. For this reason, passive thermal control is less failureprone. Lastly, this design would greatly benefit the mission by having a thermal system that does not draw power fromthe battery. This means that a smaller and lighter battery could be used for the mission.

A con of this approach is that the amount of insulation needed may make the payload too bulky and could addunneeded weight and bulk.

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Pros ConsVery simple design Potentially bulkyLess moving parts More massive

Small chance of unexpected failureViable for prolonged flight (no power involved)

Table 7. Passive Thermal Control Pros and Cons

4.2.2. Active and Passive Thermal Control

An active and passive thermal control system allows for the addition of a heater within the payload that wouldreduce the amount of insulation required to keep the internal temperature above the target temperature. This designwould still benefit from the insulation material trade study as insulation is still a vital part of the system.

One pro of this method would be that less insulator material would be needed to achieve the same result. Thiswould solve the passive thermal control system problem of making the payload too bulky and massive. Another proof this design is that it would allow for the team to choose a specific temperature to hold the payload at. This may helpin the data analysis of other sensors that benefit from being in a set temperature range.

One con of this approach is that it is the more complex design. Because it is more complex, there is additional riskfor failure through either faulty software or hardware. Another con is that the heater will draw power from the batterywhich means that either a larger battery will be needed or the flight time will need to be shorter.

Pros ConsReduces insulation material needed for passive control More prone to failure than just passive

Allows for customizable temperature within the payload Draws power (larger battery or shorter flight)

Table 8. Active and Passive Thermal Control Pros and Cons

4.3. Tether Design

The tether is critical to the function of the integrated balloon and orbiter system. The tether provides the physicalconnection between the balloon and orbiter and allows for wired communications between any electronics on theballoon (termination systems) and the primary electronics suite on the orbiter. Due to the orbiting requirements ofthe payload, the tether will need to be fairly long, which may have an adverse impact on launch operations if notproperly implemented. Three tether designs have been chosen for further consideration: a static tether of fixed-length,a passively unspooled fixed-length tether, and a dynamically unspooled winch-controlled tether. Each design will beanalyzed for impact on system transportation, launch operations, and post launch operations. System transportationfocuses on moving the balloon and payload from a storage location (likely the aerospace building) to a launch siteto begin the mission. Launch operations encompasses the logistics of launch, accounting for manpower requirementsand space for launch staging. Post launch operations addresses the operation of the orbiter attached to the balloonat all times occurring after launch through mission termination. It is critical that the tether is fully deployed withoutirregularities during launch, as this will permit the orbiter to achieve it’s maximum intended orbit for a mission.

4.3.1. Static Fixed-Length Tether

The static fixed length tether is the simplest of the three designs under consideration. This design will require atether length to be determined prior to launch, based on the given mission profile. As shown in Figure 9 the tethermay be as long as 10 m to meet the orbiting requirements of the payload with the largest anticipated balloon diameter.Given a maximum potential length of 10 m it may be stated that launch operations will require a minimum of 10 mlinear meters to prepare the vehicle for launch. In the event that the balloon is launched under high winds it may not bepossible for one person to complete the launch independently. Additionally, when transporting the system to a launchsite the tether will have to be stored, likely on a spool. Manually removing the tether from its storage device will addtime and complexity to the launch. However, this tether design has the lowest likelihood of introducing knots, snags,and tangles (i.e. proper deployment) to the tether during the launch procedure. Once launched the tether will remain atthe predetermined length and will not consume any power or resources from the systems onboard. Additionally, sincethere is no deployment during launch this style of tether is the easiest to manufacture.

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Figure 9. Static Fixed-Length Tether Deployment

Pros ConsMinimal chance of tangling Difficult to launch

Easy to manufacture Fixed length in flight

Table 9. Static Fixed-Length Tether Pros and Cons

4.3.2. Passively Unspooled Tether

The passively unspooling tether uses a fixed length tether like the static fixed-length tether, but it deploys to itsfinal length passively during the launch process. The length of the tether will be predetermined for a given missionprofile. Once the tether length has been determined the tether will be installed onto a spool deployment system. Thisdeployment system will allow for the tether to deploy automatically as the balloon rises during the launch withoutany operator management. The spool additionally will serve to transport the tether, removing a step from the launch

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procedure when compared to the static fixed-length tether. Due to the compact nature of the spool tether deployment,the launch process will be manageable by a single person and will not require a large launch site. Despite the advan-tages in launch logistics that are offered by this solution, it does present some additional risk with respect to tethertangling during deployment of the tether at launch. This risk is fairly low however, as the spool deployment methodhas decades of documented success in applications such as sewing machines and fishing pole reels. After launch thereis no opportunity to adjust the tether as it is of a predetermined and fixed length.

Figure 10. Passively Unspooled Tether Deployment

Pros ConsSmall launch footprint Potential to tangle

Simpler in construction than winch Fixed flight lengthIntegrated transport and launch spool

Easier to launch

Table 10. Passively Unspooled Tether Pros and Cons

4.3.3. Dynamically Unspooled Tether

The dynamically unspooled tether consists of a method to fully deploy the tether on command after the balloon hasbeen launched. Of the three options this is the most space efficient and the easiest to launch; the tether does not addany hazard to launch operations. After a successful balloon launch, the tether would be unspooled to the target lengththrough the use of a winching device. The use of an onboard winch also permits the tether length to be adjusted mid-flight if desired. However, the inclusion of a winch onboard adds additional mass to the system and will consume someof the limited electric power from the batteries to operate. Additionally, winch controls will need to be developed,which adds to the software and telemetry complexity. This method has a fairly low likelihood of getting tangled as thetether will remain taught due to the weight of the payload, however the addition of a winch provides another point ofmechanical system failure. The primary failure concern will be turbulence during winch deployment, as a decrease inthe force exerted on the tether may cause the winch to free spool, causing a jam. In the event the winch fails to deploy

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at all, the mission would be limited to a single data collection on descent as the orbiter would not have freedom toorbit.

Figure 11. Dynamically Unspooled Tether Deployment

Pros ConsLow tangle likelihood Adds mass and complexity

Easy to launch Uses onboard powerDynamic tether lenght during mission

Table 11. Dynamically Unspooled Tether Pros and Cons

4.4. Venting Design

In order to achieve neutral buoyancy at a desired altitude, the balloon must vent some of its helium. After ventingsome helium, the balloon must then be able to retain the rest of its helium until flight termination. This presents achallenge; if helium is leaking out of the balloon after the initial vent, then it will be significantly more difficult toundergo the desired altitude oscillations between 25 km and 35 km. This problem can be solved by implementing avalve that can properly open and close when needed. In exploring this design space, solenoid and electric motor valveswere analyzed.

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4.4.1. Solenoid Valve

Two main types of solenoid valves will be considered in this section: direct acting valves and pilot-operated valves.Solenoid valves operate by energizing an electromagnet [16]; this generates a substantial magnetic field, which pulls theflow-halting device out of the flow path. This allows venting to occur. Once the electromagnet is no longer energized,the spring returns to its usual position and the flow is halted. This configuration would need a significant power sourceto generate the electric field and also a controller to operate the solenoid when needed. This is an active valve, meaningthat knowledge of the balloon’s pressure and surrounding atmosphere is needed.

Direct acting solenoid valves are mostly used in small-size (diameters typically under 25 mm), low pressure en-vironments. This valve is able to open with zero pressure. It is faster than other solenoid valves in terms of valveconnection and disconnection. Power consumption for this valve can be relatively high, ranging from 5 to 20 W. Oneof the main problems with this design is that end-of-life venting will require a significant amount of power; it will benecessary to terminate flight when there is still enough power to sufficiently vent the balloon.

Pilot-operated solenoid valves are typically used in large diameter and high pressure conditions. Pilot pressure isrequired for this type of valve in order for it to be opened. Power consumption for this valve is low (0.1-0.2 W). Dueto the high-pressure requirement of this valve, though, it will not be considered further within the design space. Thevalve used in the HYFLITS design has a similar design to this one. Similarly, a servo spins and either opens or closesthe gap to allow air to flow or be contained respectively. The difference it that the HYFLITS has a valve that operatesvertically and is constantly drawing power to push up against the opening. In this design, it will be improved so thatthere is no waste of power.

Pros ConsReliable actuation Requires significant battery power

Valve housing is typically included Requires logic controller

Table 12. Solenoid Valve Pros and Cons

4.4.2. Electric Motor

The primary type of valve used with electric motors utilizes a linear motion. Linear electric valves are used withglobe, diaphragm, pinch, and angle valves [14]. These type of valves use a sliding stem that opens or closes a valve.Linear type actuators require tight tolerances, since any leakage will pose a large problem. As seen in Fig. 13, thevalve operates by simply pushing or pulling a valve in front of the gas flow with an electric motor. These valves arehighly reliable and can be precisely controlled. Electric motors such as servos offer the benefit of intermediate positioncontrol. This would allow a logic controller to place the valve in a partially open or partially closed state, which willbe useful during end-of-life venting.

These valves are actively controlled; this means that a logic controller with knowledge of the local conditions mustalso be placed on board. While there is a wide variety of electric motors, the varieties available to use in terms ofsize and weight have limited closing power. This could be a potential issue as higher pressures may be needed togenerate the required force. Another quality to consider is the actual valve mechanism used with an electric motor. Toutilize this configuration, a unique housing and valve design would need to be integrated with the motor. This designintroduces greater possibility for leaks to occur. This contrasts with the previous design, since valve-solenoids aretypically sold as single units with their housing.

The main benefit to using electric motor valves is that many of them require power only while they open or close;this is desirable over the solenoid valves because they require power throughout the entire duration they are actingagainst the spring.

Pros ConsAble to remain partially open Requires tight tolerances

Reliable actuation Requires logic controllerUses less power than solenoid valves Potentially weak closing force

Table 13. Electric Motor Valve Pros and Cons

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Figure 12. Diagram of Solenoid Valve

Figure 13. Diagram of Electric Motor Valve

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4.5. Balloon Selection

Balloon selection is one of the crucial elements of the HALO mission as it serves as the primary lifting source forthe payload system. This design must be decided upon early because the balloon properties will directly affect themass allowance of the payload. It is critical to consider the expansion of the balloon as it ascends to altitudes up to35 km. Only balloons which are capable of withstanding the low-temperature and low-pressure conditions of thesealtitudes will be considered. Table 14 shows a table with various parameters on different civilian grade balloons. Thistable contains information about the burst diameter, neck size and length, burst altitudes for a given payload mass, andcost of specific High-Altitude Balloons (HAB).

Balloon Neck Length Neck Size Burst Diameter Burst Altitude Range CostHAB 350 3 cm 12 cm 412 cm 14.6 - 21.3 km $30.00HAB 600 3 cm 12 cm 602 cm 22.9 - 27.4 km $45.00HAB 800 3 cm 12 cm 700 cm 27.4 - 30.5 km $65.00HAB 1000 3 cm 12 cm 786 cm 30.6 - 33.5 km $75.00HAB 1200 3 cm 12 cm 863 cm 29.3 - 33.5 km $90.00HAB 1500 3 cm 12 cm 944 cm 32.0 - 35.1 km $105.00

HAB-TX-1500 3 cm 12 cm 944 cm 32.0 - 38.1 km $120.00HAB 2000 5 cm 12 cm 1054 cm 32.6 - 35.7 km $230.00HAB 3000 5 cm 12 cm 1300 cm 35.1 - 38.1 km $360.00

Table 14. Balloon Options and Properties [12]

The naming convention for the above balloons denotes the balloon mass in grams; for example, the HAB 350has a mass of 350 g. In order to ensure mission success, it will be necessary to choose a balloon that will not burstat target altitudes. While maintaining a safety margin for burst altitudes, a balloon with the least mass and cost willbe favorable. This leads to the selection of the HAB 1000, HAB-TX-1500, and HAB 3000 to be considered for themission; other balloons were not chosen because they did not provide sufficient benefits relative to their costs. Forexample, the HAB 1200 was not chosen since it has a worse burst altitude range, higher mass, and higher cost thanthe HAB 1000. It is important to note, however, that this table is a useful reference since any of these balloons may beused when it comes to the testing phase of the mission.

Plots relating initial balloon volume as well as helium mass versus balloon diameter at altitude can be foundbelow. These plots will be useful in determining the initial helium fill and comparing estimated balloon diameterto burst diameter. It is important to note that these plots are assuming a calorically perfect gas; the ideal gas lawequation is utilized in calculating balloon volumes at varying altitudes. Further, standard atmospheric conditions areassumed. These assumptions may have significant impact on predicted balloon volume, and a sensitivity analysis willbe conducted at a later time.

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Figure 14. Initial Balloon Volume vs Diameter at Altitude

Figure 15. Helium Mass vs Diameter at Altitude

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4.5.1. HAB 1000

The HAB 1000 balloon has a mass of 1,000 g, which is the lowest mass balloon to be considered. This is ideal interms of weight; less lifting force will be required to control altitude if the balloon is of a lower mass. Its neck lengthis 3 cm, which will be significant when considering venting and termination options. Its burst range is from 30.6 km to33.5 km, which would be acceptable to reach levels of success 1 and 2. Finally, its cost is $75.00, which is the lowestof the considered balloons.

Pros ConsLowest mass Lowest burst altitudeLowest cost

Table 15. HAB 1000 Pros and Cons

4.5.2. HAB-TX-1500

The HAB-TX-1500 balloon has a mass of 1,500 g, which is the second lowest balloon mass to be considered. Itsneck length is also 3 cm, which will be relevant when considering venting and termination options. Its burst range isfrom 32.0 km to 38.1 km, which would be suitable for reaching the third level of success. Finally, its cost is $120.00,which is much less expensive than the more massive balloons.

Pros ConsRelatively low mass Wide burst altitude range

High maximum burst altitudeRelatively low cost

Table 16. HAB-TX-1500 Pros and Cons

4.5.3. HAB 3000

The HAB 3000 balloon has a mass of 3,000 g, which is the most massive balloon to be considered. Its neck lengthis 5 cm, which will be relevant when considering venting and termination options. Its burst range is from 35.1 km to38.1 km, which would be ideal for reaching the third level of success. Finally, its cost is $360.00, which is much moreexpensive than the other balloon options.

Pros ConsHigh minimum burst altitude Highest massHigh maximum burst altitude Highest cost

Table 17. HAB 3000 Pros and Cons

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5. Trade Study Process and Results5.1. Lifting Orbiter Design Trade Study

After the balloon obtains neutral buoyancy and begins to rise in steady state conditions at approximately 2 ms ,

the vertical forces will cancel to zero. This will force the balloon to have zero acceleration and maintain a constantvelocity. Two main vertical forces that are in play: drag on the balloon and payload, as well as the lift from the orbiter.When comparing the drag by the balloon with the drag of the payload, the payload’s vertical drag can be neglectedsince it will be very small relative to the balloon drag. This leaves the drag on the balloon which is calculated byEquation 6 below.

D =12ρV2S CD (6)

Using a density at 30 km as 0.017 kgm3 , the maximum speed defined by requirements as 2 m

s , and using the surfacearea of a hemisphere for the cross sectional area of 78.74m2, the drag can be calculated. This cross sectional area wasobtained by approximating the maximum balloon diameter at altitude with the ideal gas law. The coefficient of dragwas found from similar weather balloons to be around 0.285. This brings the drag force to 0.763 N. In other words,0.763 N is the minimum lifting force required to maintain a constant speed of 2 m

s in ascent. Equation 7 displays theequation which calculates the minimum airfoil area required to produce this lifting force. It was obtained by settingthe lift from Equation 1 equal to the drag from Equation 6.

S =1.526[N]ρV2CL

(7)

The density value is the same, but the speed will be the value that the wing is traveling at. This is constrained bya maximum speed of 10 m

s which is used to determine the maximum lift. The coefficient of lift used is approximatedat 1.3 which is a reasonable initial guess until a certain airfoil is chosen. Using these values, the area of wing neededwould be about 0.69 m2. Considering that the area is the span of the wing multiplied by the chord length, this resultsin a giant wing compared to the small payload.

5.1.1. Metrics

In this section, as well as all other metrics sections, the scoring system will be based on potential levels of success.Any metric with a score of 5 will be ideal for achieving the highest level of success, level 3. Any metric with a scoreof 4 will be capable of reaching the highest level of success, but ideal for the 2nd level of success. A score of 3 willbe capable of reaching the 2nd level of success, but ideal for the 1st level of success. A score of 2 will be capableof reaching the 1st level of success, but not ideally. A score of 1 indicates a design option which will likely not bechosen, as it would require significant trade-offs in order to be suitably implemented.

This is not a perfect scoring system by any means, but it will serve as a baseline for determining the true meaningof an option’s score. It is especially difficult to rate qualitative metrics on this scale, though an attempt will be made.For the lifting orbiter design trade study, the following metrics were deemed relevant: control system complexity,mass, power consumption, and cost.

Control System Complexity: Control system complexity is an integral factor when considering the best liftingorbiter design. The Wing with Mounted Propeller and Wing on a Rotational Tether designs both require an advanced,effective control system; they must be maintaining specific orbiter angles of attack while remaining stable under windperturbations. If anything goes wrong with these control systems and the payload does not properly orbit, then no liftwill be generated and the mission will be a failure. The third design, however, does not require a very advanced controlsystem because the orbiter is not the primary lifting force. Even if the orbiter fails, the balloon will still be able to liftitself to target altitudes. For these reasons, the control system complexity parameter will earn a weight of 0.1.

Requires a Complex Control System? No YesScore 5 3

Table 18. Control System Complexity Scoring

Mass: Mass is a significant parameter for almost every aspect of this mission, since a higher overall mass willrequire a higher lifting force to effectively control the altitude. Considering that weight limitations will need to be met,

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as set forth by FR.1, this becomes a pivotal system for weight constraints for all other systems. For this reason, themass parameter has earned a weight of 0.3. The mass of each system will be predicted using typical masses of motorsand propellers, as well as an approximated mass of the payload.

The first two designs will have very similar masses; they both require a large airfoil, a motor, and servos to actuatechanges in angle of attack. While the third design does include two motors, it will not require a very large wing tooperate.

Mass [kg] 0.00-0.60 0.61-0.90 0.91-1.20 1.21-1.5 1.5+Score 5 4 3 2 1

Table 19. Mass Scoring

Power Consumption: Power consumption is one of the more critical parameters for the lifting orbiter system.Power is required not only for data collection and transmission, but also for driving any propellers and motors. Thissystem is expected to be the main power consumer of all systems within the mission. For this reason, power consump-tion earns a weight of 0.4 for this design space. While this is a necessary parameter to quantify when evaluating theproposed lifting orbiter designs, this proves difficult to measure at this early stage. In light of this, general approxima-tions were made as to the power consumption of each design.

The first design should theoretically require the least power, due to its nature of including only a single propelleron the airfoil. The second design should require the median amount of power, since a motor is driving the circularmotion of the tethered-orbiter entirely. The third design will certainly require the most power, as a more powerfulmotor is implemented for the purpose of generating the desired lift for the balloon system.

Power Consumption [W] 0.0-4.0 4.1-8.0 8.1-12.0 12.1-16.0 16.1-20.0Score 5 4 3 2 1

Table 20. Power Consumption Scoring

Cost: Cost is also a significant parameter for almost every aspect of this mission, since it is desired by thecustomer that a single launch will cost no more than $2,000. Further, development and testing for the entire missionmust cost no more than $5,000. This system is the main weight constraint of the payload, and as a result will largely bythe main consumer of cost. As a result, this metric earns a weight of 0.2. The HYFLITS balloon and sensor packagecost approximately $1,000 per launch. This means that approximately $1,000 more can be spent in developing the restof our systems; the cost scoring stems from this budget constraint.

It is estimated that the first two designs will fall into a score of 4, as they require most of the same parts. Theirestimated cost will be somewhere in the $401-$600 range. The third design is estimated to be slightly less expansivedespite the addition of a second motor, since the wing will be of a significantly smaller size.

Cost [$] 0-400 401-600 601-800 801-1000 1000-1200Score 5 4 3 2 1

Table 21. Cost Scoring

5.1.2. Scoring

Based on the above metrics, a lifting orbiter design trade study matrix was developed and analyzed. The SeparateOrbiter and Altitude Control system scored the highest with a 3.95; it is predicted to have a simple control system,low mass, low cost, and relatively low power consumption. The score of 3.95 suggests that level 3 success should beachievable. Both of the other two designs trailed only by a small margin which means that if any complications arisewith the separate orbiter and altitude control system, these other designs may be reconsidered.

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Category Weight Wingwith Mounted Propeller

Lifting Wing on aRotational Tether

Separate Orbiterand Altitude Control Motors

Control SystemComplexity 0.10 3 3 5

Mass 0.30 4 4 4.5Cost 0.20 4 4 4.5

Power Consumption 0.40 4 3.5 3Score 1.00 3.9 3.7 3.95

Table 22. Lifting Orbiter Design Trade Study Matrix

5.2. Thermal Control Design Trade Study

5.2.1. Metrics

The purpose of this thermal trade study is to compare the types of foam materials available for insulation. In orderto perform a study on the effectiveness of each insulation, the model described in section 4 was used. This is anarbitrary heat generation model used for the sole purpose of comparing the insulation materials. This trade study willnot be used to choose between a purely passive system or active and passive combination thermal system. The reasonfor this is that there are too many unknowns at this time because the exact design of the orbiter and altitude systemsare currently unknown as well as their heat generation capability. For this reason, this trade study does not examineactive thermal control.

Four main metrics will be considered within the thermal selection trade study: thermal conductivity, volume (basedon model), mass (based on model), and the cost (based on the model) of the foam materials. With this in mind, eachcategory was given a specific weighting, in order to prioritize the most important parameters that would provide thebest possible choice.

Three main foam materials were considered in this trade study which include: polyurethane foam, extrudedpolysyrene (XPS), and phenolic foam. The thermal conductivities, costs per square centimeter, and densities of thesematerials were researched and then plugged into the thermal model (Appendix A) to find the thickness, total mass, andtotal cost associated with each material.

Thermal Conductivity: Thermal conductivity is a very significant metric when it comes to the choice of insu-lation material. This is such an important metric because it inversely relates to the thickness of insulation required tolimit the heat transfer into the payload down to a workable level. In other words, a lower thermal conductivity willultimately result in a lower thickness of insulation required. This in turn results in the insulation being less bulky.Because the orbiter payload will need to also be a functional aerodynamic body, reducing the bulk of the insulationcould be critical to mission success. For this reason, the thermal conductivity metric earns a weight of 0.2. This willbe further represented in the thickness metric below.

Thermal Conductivity [ WmK ] 0.020-0.023 0.024-0.026 0.027-0.029 0.030-0.032 0.033-0.035

Score 5 4 3 2 1

Table 23. Thermal Conductivity Scoring

Thickness (from Model): The thickness metric is significant when considering which insulation material tochoose because it directly relates to the bulk of the system. The bulk of the system is important because the orbitersystem will be an aerodynamic body that will need to be sleek and low-mass. For this reason, the thickness parameterearns a weight of 0.35. The following thickness metrics were calculated using the model described above.

Thickness [cm] 2.0-2.4 2.5-2.8 2.9-3.2 3.3-3.6 3.7-4.0Score 5 4 3 2 1

Table 24. Thickness Scoring

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Total Mass (from Model): The mass metric is significant when considering which insulation material to choosebecause the mission has a strict mass constraint. For this reason, mass is arguably one of the most important aspects ofthe chosen insulation material with the lowest values being preferred. For this reason, the mass metric earns a weightof 0.225. The following insulation mass metrics were calculated using the model described above.

Total Mass [g] 50-60 61-70 71-80 81-90 91-100Score 5 4 3 2 1

Table 25. Total Mass Scoring

Total Cost (from Model): The cost metric is significant when considering which insulation material to choosebecause the mission has a strict cost constraint. For this reason, the cost metric earns a score of 0.225. The missionstatement requires that the overall payload construction remain under $2,000 per launch. The following insulationcost metrics were calculated using the model described above. This category is unique in that it incorporates a scorefrom 1-8. This is due to the fact that certain materials are signi f icantly more expansive than others, and this has to beheavily reflected in the scoring.

Total Cost [$] 0-10 10-20 20-30 ... 70-80Score 8 7 6 ... 1

Table 26. Total Cost Scoring

5.2.2. Scoring

The weighting of each thermal category was decided based upon the importance of the parameter. The mostimportant parameter was decided to be thickness. This parameter is important because it prevents the insulationfrom getting too bulky and cumbersome on the payload. For that reason it was given a weight of 0.35. The thermalconductivity was given a weight of 0.20 because it is directly correlated to thickness but still important for preventingheat transfer into the payload. The remaining two parameters are total mass and cost. The total mass and cost weredetermined to be equally important because the mission is severely limited by both mass and cost.

Based on the results of the trade study illustrated in Table 27, polyurethane was determined to be the best foammaterial to use as insulation for the payload. Polyurethane scored very well in all technical categories, and is veryinexpensive. This makes it the optimal insulator for this thermal design space.

Category Weight Polyurethane Extruded Polystyrene Phenolic FoamThermal Conductivity 0.20 4 1 5

Thickness 0.35 4 1 5Total Mass 0.225 3 1 5Total Cost 0.225 8 7 1

Score 1.00 4.67 2.35 4.10

Table 27. Insulator Material Trade Study

5.3. Tether Design Trade Study

5.3.1. Metrics

The tether metrics analysis is purely qualitative, due to the subjective nature of the tether trade study. While therewill be numerical analysis at a later date to select an appropriate material, the most challenging design choices thataccompany the tether are currently non-numerical. A variety of commonplace processes are used as stand-ins for ex-pected performance, and all three designs present the opportunity for rapid prototyping, testing, and refinement furtherin the design process.

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Ease of Launch: Per customer requirement the balloon and orbiter system must be easy to launch. Ideally launchoperations will require no more than one person, but may require two launch operators depending on the chosen de-sign. The primary contributor to the number of launch operators required for a standard launch is the size of the launcharea. If a given design permits launch entirely from a man-portable launch platform it can be operated by a single op-erator. However, if a tether needs to be extended (fully or partially) additional launch operators may be required tofacilitate the launch, especially during inclement weather. For these reasons, the ease of launch metric earns a weightof 0.5. This is the most significant metric for the tether design space. The five scoring categories for ease of launchare outlined below.

Two-Phase Launch - 5: The two-phase launch is the highest score possible in this category. To meet this scorerequirement the launch may be completed from a self contained launch platform by a single operator in two distinctphases. The first phase consists of installing the balloon to the launch platform and filling the balloon to the desiredvolume. Phase two releases the balloon, after triggering the release the launch operator is has completed the launch.

Two-Phase+ Launch - 4: The two-phase+ launch is identical to the two-phase launch except it requires the operator totrigger the deployment of the tether after balloon release. This is primarily a concern for the active winch deploymentmethod, but depending on the final design of the launch platform it could also describe the passively unspooling launch.

One+ Operator Launch - 3: This launch can be easily completed by a single operator under standard launch condi-tions. Standard launch conditions require low winds and a large launch site. In the event of launch during inclementweather or in an enclosed launch site another launch operator may be added to help ensure mission success.

Two Operator Launch - 2: The two operator launch will require two operators to ensure mission success regardlessof launch conditions.

Three Operator Launch - 1: This launch will require three or more launch operators to successfully launch the balloonand payload under all circumstances. This is in clear violation of customer requirements and will be avoided at allcosts.

Ease of Launch Two-Phase Two-Phase+ One+ Operator Two Operator Three OperatorScore 5 4 3 2 1

Table 28. Ease of Launch Scoring

Mass: Due to the limited mass budget, it is critical to minimize dead weight on both the balloon and payload. Anyadditional weight will require more helium to lift to the target altitude and will require controlled ascent subsystem tolift more mass, which will slow the ascent and consume more power. For this reason, the mass metric earns a weightof 0.2 for this design space. The categories are divided by mass added to the system, since both the static tether andpassively unspooled tether require no additional mass on the system, they have achieved the maximum score. Standardhobby-sized winches weigh anywhere from 50 g to 100 g. In this preliminary state, it is assumed a typical hobby winchwith a mass of approximately 75 g is to be used for the active unspooling launch.

Airborne Mass [g] 0-24 25-49 50-74 75-99 100+Score 5 4 3 2 1


Tangle Likelihood: A tangled tether has the potential to ruin the mission by limiting the payload to a reducedorbit. While this is a very important parameter, it is closely related with ease of launch. If launch goes smoothly, therewill be a lower likelihood of tangled tethers. For this reason and because ease of launch is already weighted heavily,the tangle likelihood metric earns a weight of 0.1. Tether designs are ranked based on the likelihood of a tangle duringlaunch and deployment. The following is used to describe the chance of tangles for a given design:

No Tangles - 5: To achieve a maximum score in the tangling category there will be no chance of tangling the tether

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during launch or deployment. This requires a manual deployment of the tether on the ground by a launch operator.

Low Likelihood Tangles (Fixable) - 4: Tangles are permitted so long as there is a low chance that there will be atangle, if a tangle occurs it will be fixable by the launch operator.

High Likelihood Tangles (Fixable) - 3: Tangles are permitted. There is a high chance that there will be a tangle,though if a tangle occurs it will be fixable by the launch operator. This is preferred over the low likelihood non-fixableas it delays launch but does not impact mission success.

Low Likelihood Tangles (Non-Fixable) - 2: Tangles are permitted so long as there is a low chance that there will be atangle, if a tangle occurs it will not be fixable by the launch operator and may impact mission success.

High Likelihood Tangles (Non-Fixable) - 1: Tangles are permitted so long but there is a high chance that there will bea tangle, if a tangle occurs it will not be fixable by the launch operator and may impact mission success.

Tangle Likelihood No Tangles Low (Fixable) High (Fixable) Low (Non-Fixable) High (Non-Fixable)Score 5 4 3 2 1

Table 30. Tangle Likelihood Scoring

Transportability: Transportability reflects ease of transport and the transition from transport to launch. Trans-portation systems that also function in a launch capacity are preferred over those that require uninstallation from atransportation system prior to launch. Furthermore, systems that must be uninstalled and reinstalled onto anotherlaunch device are less preferred that those that require only uninstallation. For these reasons, the transportability met-ric earns a weight of 0.2.

Transportation is Launch System - 5: The tether will be installed into the launch system and transported in that con-figuration. The tether is ready to be launched with no additional steps.

Transportation is Launch Compatible - 4: The tether will be installed into a separate component of the launch system.The separate component will need to be integrated into the launch system prior to launch.

Uninstallation Only - 3: The tether will need to be uninstalled from the transportation system to launch. Once re-moved from the transportation system it is ready to launch.

Direct Reinstallation to Launch System - 2: The tether will be transported independently from the launch system.Prior to launch the tether will be uninstalled from the transportation system and reinstalled directly onto the launchsystem

Indirect Reinstallation to Launch System - 1: Tangles are permitted so long as there is a low chance that there will bea tangle, if a tangle occurs it will not be fixable by the launch operator and may impact mission success.

Transportability Launch System Launch Compatible Uninstallation Direct Reinstallation Indirect ReinstallationScore 5 4 3 2 1

Table 31. Transportability Scoring

5.3.2. Scoring

Considering the metrics outlined in the previous sections, a trade study matrix was created and analyzed. It wasfound that the statically unspooling tether is the best of the considered designs. The combination of high transporta-bility, low likelihood of fixable tangles, no additional mass, and single operator launch represent the best combinationof desired performance characteristics.

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Category Weight Static Tether Passively Unspooled Tether Dynamically Unspooled TetherEase of Launch 0.50 1.5 4.5 4.0

Mass 0.20 5.0 5.0 4.0Tangle Likelihood 0.10 5.0 4.0 2.0Transportability 0.20 3.0 5.0 4.0

Score 1.00 2.85 4.65 3.80

Table 32. Tether Design Trade Study Matrix

5.4. Venting Control Trade Study

5.4.1. Metrics

In considering the venting control trade study, the following metrics were investigated: mass, power, cost, andreliability. These metrics will be explained and analyzed in the following sections.

Mass: Mass is a critical element for every trade study, since any excess mass will make lift generation significantlymore difficult. The mass of solenoid valves and of electric motor valves were investigated independently. While thesevalves come in a very wide range of sizes, typical masses of valves that may be feasible in this design were researchextensively. It was found that solenoid valves have a typical mass of around 0.1 kg, and electric motor valves have atypical mass of around 0.3 kg. These are rough estimates, and will be narrowed down at a future date. Since thesemasses are relatively low compared to the payload and balloon mass, the weight for this metric will be only 0.15;metrics such as power, cost, and reliability will be much more significant for this design.

Mass [kg] 0.00-0.15 0.16-0.30 0.31-0.45 0.46-0.60 0.61-0.75Score 5 4 3 2 1


Power Consumption: The operational lifespan of the balloon is dictated in part by power from the batteries.Therefore, reducing power consumption will extend the operating time of the balloon system. A valve that usesmore power will require a more powerful battery, which will increase weight. Reducing power usage needed for thevalve would allow more power to be used in the orbiting system and sensing devices. Power will be one of the moresignificant parameters when it comes to venting decisions, and as such earns a weight of 0.3.

The main difference when it comes to the two valve types, in terms of power, is that the solenoid valve requirespower throughout the entire venting process. The electric motor valves, on the other hand, require power only whileshifting the flow-halting device; in other words, the electric motor valves require power over much smaller timeduration than the solenoid valves. In order to determine how much more total power the solenoid valve requires, itwould be necessary to know the amount of time it takes for the balloon to vent to neutral buoyancy. For this reason,power consumption will be scored quantitatively based off of the typical wattage of each option.

Power Consumption [W] 0.0-2.1 2.1-4.0 4.1-6.0 6.1-8.0 8.1-10.0Score 5 4 3 2 1

Table 34. Power Scoring

Cost: Cost is a significant metric for the system since the project is aiming to minimize the cost per launch; percustomer request, the cost per launch should be at most $2,000. Valves vary significantly in cost based on quality ofconstruction and complexity. Cost is of roughly the same significance as the mass and power consumption metrics.For this reason, the cost metric for this design space earns a weight of 0.2.

While each type of valve comes in very wide price ranges, it was generally found that solenoid valves are slightlyless expensive than similar electric motor valves.

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Cost [$] 0-10 10-20 20-30 30-40 40+Score 5 4 3 2 1


Seal Reliability: Reliability of the venting system does not relate the lifespan of the valve, but rather how muchconfidence can be instilled in it to function without significant leakage. Since leakage will make it significantly moredifficult to achieve desired altitudes, this metric earns a weight of 0.4.

Many manufacturers produce solenoid valves that would easily be implemented into the balloon system. Electricmotors, on the other hand, would need to be integrated into a valve system which could add larger uncertainty towardsthe confidence in the seal.

Valve Leakage Likelihood Highly unlikely Unlikely Uncertain Likely Highly LikelyScore 5 4 3 2 1

Table 36. Seal Reliability Scoring

5.4.2. Scoring

In considering the metrics outlined in the previous sections, a trade study matrix was developed in order to analyzethe optimal venting system. It was found that a solenoid valve would be best suited for this mission due to its lowmass, low cost, and high seal reliability. The electric motor, however, trailed by only 0.3 points. This suggests that ifany complications arise while developing the solenoid valve, the electric motor may be reconsidered.

Category Weight Solenoid Valve Electric MotorMass 0.20 5 4

Power Consumption 0.20 2.5 5Cost 0.20 4 3

Seal Reliability 0.40 5 4Score 1.00 4.3 4.0

Table 37. Venting Options Trade Study Matrix

5.5. Balloon Selection Trade Study

5.5.1. Metrics

Four metrics will be considered within the balloon selection trade study: minimum estimated burst altitude, maxi-mum estimated burst altitude, mass, and cost. With this in mind each category was given a specific weighting, in orderto prioritize the most important parameters that would give the group the best possible choice.

Minimum Burst Altitude: The minimum burst altitude is the most significant metric when it comes to the balloonchoice. This is because it is necessary to choose a balloon which will not burst at target altitudes; if the balloon bursts,then data cannot be collected and the mission is a failure. For these reasons, a weight of 0.35 is assigned to this metric.Based on the following table, it will be critical to choose a balloon with a minimum burst altitude score of at least 2for level 1 success, 3 for level 2 success, and 4-5 for level 3 success. These minimum burst altitudes were listed on theballoon provider’s website, and were tested using varying payload masses of approximately 1 kg.

Minimum Burst Altitude [km] 35-40 30-35 25-30 20-25 15-20Score 5 4 3 2 1

Table 38. Minimum Burst Altitude Scoring

Maximum Burst Altitude: The maximum burst altitude is another significant parameter when it comes todeciding which balloon to use. It is ideal to have a maximum burst altitude that is significantly higher than the actual

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target altitude; this would suggest that the balloon is capable of going even higher without bursting. For this reason,this metric was given a weight of 0.3. The following table depicts the scoring definition for this parameter. It will becritical to choose a balloon with a maximum burst altitude score of at least 2 for level 1 success, 3 for level 2 success,and 4-5 for level 3 success. These maximum burst altitude were listen on the balloon provider’s website, and weretested using a payload with mass of approximately 3 kg.

Maximum Burst Altitude [km] 35-40 30-35 25-30 20-25 15-20Score 5 4 3 2 1

Table 39. Maximum Burst Altitude Scoring

Mass: The mass is a significant parameter when it comes to balloon choice because it directly effects the down-ward force acting on the balloon system. A heavier balloon will require a greater lift force to increase its altitude,which will require more force to be generated. It is difficult to generate large amounts of lifting force in the low-pressure and low-density environment of the stratosphere, and so it is desirable to require as little lift as possible. Theminimum mass listed is only 1,000 g, since lighter balloons are not capable of maintaining altitudes required by thisproject. With this in mind this parameter was given a weight of 0.15.

Mass [kg] 1.0-1.4 1.4-1.8 1.8-2.2 2.2-2.6 2.6-3.0Score 5 4 3 2 1


Cost: The cost of the chosen balloon is a significant parameter due to the single-use nature of the mission; thepayload and balloon will not be recovered. Per customer request, the total cost per launch shall be no more than$2,000. For this reason, this metric was given a weight of 0.2. The scoring table for the cost metric can be foundbelow.

Cost [$] 61-120 121-180 181-240 241-300 301-360Score 5 4 3 2 1


5.5.2. Scoring

Using the above tables from the metrics section, (tables 38 - 41) the following trade study matrix was created andanalyzed. Table 42 displays that the HAB-TX-1500 balloon will best fit the mission requirements needs in terms offacilitating the desired altitudes, as well as minimizing mass and cost. The HAB 1000 trails by only 0.15 points, andwill be considered if the total mass of the payload is sufficiently low.

Category Weight HAB 1000 HAB-TX-1500 HAB 3000Min Burst Altitude 0.35 4 4 5Max Burst Altitude 0.30 4 5 5

3 Mass 0.15 5 4 1Cost 0.20 5 5 1

Total Score 1.00 4.35 4.5 3.6

Table 42. Balloon Selection Trade Study Matrix

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6. Selection of Baseline Design6.1. Lifting Orbiter Design Selection

Based on Table 22, the best scored design was the separate orbiter and altitude control system. This design (designin 4.1.3) scored the best in the categories of control system complexity, mass, and cost. Since there was a very smallmargin between this design and the two trailing designs, this choice is not fully determined and may be reconsideredupon further investigation. This system will consist of one motor for the pupose of generating lift for the balloon, andone motor for the purpose of generating the required rotational motion for the payload.

6.2. Thermal Options Selection

Based on preliminary evaluation, the thermal system selected will be a combination of passive and active control.The exact heater that will be used in the active thermal control subsystem is currently unknown and will remainunknown until the heat dissipation of essential electronics operating in the payload are measured.

A trade study was done on possible foam materials that could be used as passive insulation for the thermal controlsystem. Based on the trade study completed in section five of this report, the best foam material for this mission ispolyurethane. Polyurethane offered the best performance (small thickness) at a reasonable mass and cost compared toextruded polystyrene and phenolic foam. Extruded polystyrene had too high of a thermal conductivity that resultedin a large thickness that would not be viable for the orbiter designs. Phenolic foam had very low thickness and mass,but was too expensive to be a valid choice given that the payload is expendable. For these reasons, polyurethane is thebest thermal insulation option.

6.3. Tether Design Selection

The chosen tether design is the passively unspooling tether design. This design adds no additional weight to theballoon or payload and permits for simple transportation and launch. A properly designed launch platform will allowfor the tether to be installed onto the deployment spool for transportation and will require no additional work priorto launch. While snags are a possibility, they will be fixable by a launch operator as the tether is deploying from theground. However, the passive unspooling method is modeled after a spinningspool or spincast fishing reel, whichis known for reliably deploying lightweight lures without snagging. While launching a high altitude balloon is notthe same as casting a fishing reel, the mechanics are much the same and the fishing reel serves to provide decades ofanecdotal evidence of success.

6.4. Venting Options Selection

Through the preliminary evaluation of the different valve choices, the selected design is a solenoid valve. Theexact solenoid valve to be used is still unknown. However, there are many off-the-shelf options readily available.

The venting options trade study indicates that the solenoid valve performs well under many relevant categories.Solenoid valves can be found that are very lightweight and reliable. In addition, they can be found for relatively cheapand do not use an excessive amount of power. The solenoid valve outscored the electric motor valve in the trade studyby a significant margin.

6.5. Balloon Selection

After carefully evaluating significant balloon parameters such as minimum burst altitude, maximum burst altitude,mass, and cost, it has been determined that the HAB-TX-1500 balloon will be selected for this mission. The tradestudy matrix in Table 42 was used in making this decision; the HAB-TX-1500 scored an impressive 4.5 points, theHAB 1000 was a close second with 4.35 points, and the HAB 3000 scored only 3.6 points. The HAB-TX-1500 isideal because it will be able to reach target altitudes for level 3 success while also remaining a relatively low-mass andlow-cost selection.

6.6. Baseline Design Selection

With all of these choices considered, a baseline design is formed. A balloon consisting of the HAB-TX-1500(1500 g) balloon, carrying a wing payload in an orbiting fashion around the outer radius of the balloon with separateorbital and altitude control systems. The balloon tether will be unraveled passively during the launch/initial ascentphase of the balloon life. The payload will contain both a passive and active thermal control system, where-by passive

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insulation will be polyurethane material. The venting system will be controlled by a solenoid valve. These systemssuccessfully satisfy functional and design requirements. Systems that further need to be explored include the CDAS,necessary for safe descent speeds. Specific instrumentation will need to be explored as well for many of these systems,especially pertaining to the lifting orbiter.

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Appendix A: Design ModelingOrbiter Modeling

To begin modeling a design with a wing, the assumption is made that the balloon has achieved neutral buoyancyand is then rising in steady state at 2 m

s . The drag on the balloon can be estimated by Equation 8 below. In this case,V would be the velocity of the balloon as it is rising (2 m

s ), S is the cross-sectional area of the balloon at altitude, andCD is the coefficient of drag. This value can be estimated from the NASA [4] plot below given a Reynolds’s number(equation 9).

D =12ρV2S CD (8)

Re =ρVlµ

(9)

Using atmospheric tables to find the needed variables at altitude, a Reynolds’s number of about 16,000 can becalculated. This is very low due to the slow speed at which the balloon travels. From this, a rough estimate of a CD pf0.5 can be obtained.

Figure 16. It is assumed that the balloon will be spherical at altitude

Using this in Equation 8, the drag on the balloon can be found to be about 0.67N. At steady state, the balloon willhave a net vertical force of 0, so the lift from the airfoil must counter the drag. Setting L = D, the area of the neededairfoil can be found; using the maximum speed allowed without disturbing the data (assumed to be roughly 10 m

s untilfurther analysis) this area is about 0.6m2 which results in a very large wing. Now, the drag of the wing may need to beaccounted for since it may not be negligible compared to the balloon. This will increase the drag and thus increase thesize of the wing even more.

The radial force required to maintain a circular orbit will be approximated by the following equation (Equation 10.

Fcentripetal =mV2

r(10)

This can be related to the horizontal component of the tether force to balance the forces. θ is taken as the anglethat the tether sweeps out between the vertical (straight down) configuration and moving up towards the horizontal.From the horizontal force balance, the speed that the orbiter must travel at to retain circular motion can be related to

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the angle at which the tether swings (Equation 11). r is the radius of the circle traced out (radius of the balloon) and Tis the tension in the rope below which can be solved for by vertical force analysis (T =

mgcos(θ) ).

ΣFx = T sin (θ) =mV2

R(11)

So given this equation, to get the most lift out of the wing, the speed must be at the maximum allowed, which is10 m

s . This would result in a 70° angle from the vertical. Obviously, this model is not 100% accurate since the wakemay not be well modeled as a vertical column right below the balloon. This would mean that r would not just be theradius of the balloon, but it would be greater to get out of the larger wake. Additionally, wind gusts will also changethe force balance so the net force won’t be zero.

As far as propeller modeling is concerned, Equation 12 is the one that will be used where T is the thrust that thepropeller produces.

T =πD2

4ρ(V∞ +

∆V2

)∆V (12)

Here, there are three unknowns: the propeller thrust, propeller circle diameter, and the change in velocity acrossthe propeller blades. The objective is to find a way to pick a certain propeller/motor combination to meet the thrustrequirements. So taking design 3 in steady state as an example, the propeller must equal the drag of the balloonwhen traveling at the 2 m

s to continue traveling at that constant speed (it is assumed that the drag of the payload isinsignificant compared to the drag of the balloon given the size comparison). Taking the thrust of the propeller equalto that drag (0.67N given already in section 4), a propeller then must be chosen to find the change in velocity of thefluid. Assuming that the blade circle diameter (D) is about a foot, or 0.3 meters, the change in velocity can then besolved for using the quadratic equation below.

∆V + 2V∞∆V −(2)(4)(T )πD2ρ

= 0 (13)

Once ∆V is found, this can be added to the free-stream velocity and used in Equation 14 to find the mass flowthrough the blades assuming the density doesn’t change much.

m = ρAV (14)

Now, knowing the mass flow rate, we can convert this into a volumetric flow rate given the density. The volumetricflow rate can be thought of as a column of air with the diameter of the fan flowing through a length h. If the flow rateis higher, h will be higher as well (see Equation 15). Since the volumetric flow rate is known, h can also be solved forto get a value in m

s to see how far the column can span given a time. This unit can also be converted to inmin .

V =πD2

4h (15)

When the pitch of a propeller is given, it can be thought of as a certain diameter column of air flowing past the aira length equal to the pitch in one revolution. The blades work like a screw and allow that much through. So lastly,by dividing h by the pitch (using correct units), this will result in a certain amount of rotations per minute. h dividedby the pitch (P) would have consistent units in

min ∗1revPin . Once the RPM variable is solved for, it can be used to find a

motor that can be combined with the specified blade to obtain the thrust required. In the case of design 3, using a bladediameter of 0.3m, and a 6” pitch, we would need a motor to run at about 9,700 RPM.

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Thermal Modeling

1 %% This is the Thermal Modeling Program that will be used for the HALO Mission2 clear; close all; clc;3

4 %% Generic Values5 Ta = -60 + 273.15; % Kelvin6 Ti = -20 + 273.15; % Kelvin7 L = 20/100; % m8 ri = 5/100; % m9 Qgen = 3; % W

10

11 %% Insulation Specific Values12 % Material 1 is polyurethane13 k1 = 0.025; % [W/mK]14 density1 = 0.035; % [g/cmˆ3]15 cost_vol1 = 0.00256; % [$/cmˆ3]16

17 % Material 2 is extruded polysterene18 k2 = 0.034; % [W/mK]19 density2 = 0.030; % [g/cmˆ3]20 cost_vol2 = 0.00424; % [$/cmˆ3]21

22 % Material 3 is phenolic foam23 k3 = 0.020; % [W/mK]24 density3 = 0.035; % [g/cmˆ3]25 cost_vol3 = 0.04996; % [$/cmˆ3]26

27 %% Call Thickness Function using specific material properties28 for ii=1:329 switch ii30 case 131 k = k1;32 rho = density1;33 cost_vol = cost_vol1;34 case 235 k = k2;36 rho = density2;37 cost_vol = cost_vol2;38 case 339 k = k3;40 rho = density3;41 cost_vol = cost_vol3;42 end43 [thickness(ii),mass(ii),totCost(ii)] = GetThicknessMassCost(k,Ta,Ti,L,ri,Qgen,rho,cost_vol);44 end

1 function [thickness,mass,totCost] = GetThicknessMassCost(k,Ta,Ti,L,ri,Qgen,rho,cost_vol)2 %% Solve for thickness3 thickness = ri*(exp(k*2*pi*L*(Ti-Ta)/Qgen) - 1)*100; % [cm]4

5

6

7 %% Solve for mass8 L = L*100; % [cm]9 ri = ri*100; % [cm]

10 Ai = pi*riˆ2; % [cmˆ2]11 Ao = pi*(ri+thickness)ˆ2; % [cmˆ2]12 vol = L*(Ao-Ai); % [cmˆ3]13 mass = rho*vol; % [g]14

15

16 %% Solve for total cost17 totCost = cost_vol*vol; % [$]18

19 end

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References[1] Chapter 7. Airplane Aerodynamics and Performance, by Chuan-Tau Edward. Lan and Jan Roskam, DARcorpora-

tion, 1997, pp. 265 - 330.

[2] Battery University, ’BU-502: Discharging at High and Low Temperatures’, Retrieved August 29,2019, from https://batteryuniversity.com/index.php/learn/article/discharging_at_high_and_low_temperatures

[3] ”Controlled Weather Balloon Ascents and Descents for Atmospheric Research and Climate Monitoring”, Atmo-spheric Measurement Technology, March 7, 2016, Retrieved September 9, 2019, from https://www.ncbi.nlm.nih.gov/pmc/articles/PMC5734649/ [Accessed 09 11. 2019]

[4] Drag of a Sphere. NASA, NASA, www.grc.nasa.gov/www/k-12/airplane/dragsphere.html.

[5] Engineering ToolBox, (2003). U.S. Standard Atmosphere. [online] Available at: https://www.engineeringtoolbox.com/standard-atmosphere-d_604.html [Accessed 09 09. 2019].

[6] FAA Regulations for Kites/Balloons, from http://www.chem.hawaii.edu/uham/part101.html

[7] Grainger, ’Open Cell, Foam Sheet, Polyurethane’, Retrieved August 29, 2019, from https://www.grainger.com/product/5GCU4?gclid=Cj0KCQjww7HsBRDkARIsAARsIT45Lm5vDpt0PM_np7cVeeRBe4-N8p7qvgsJE864gIaMXGkcWKOAQCAaAoyNEALw_wcB&cm_mmc=PPC:+Google+PLA&ef_id=Cj0KCQjww7HsBRDkARIsAARsIT45Lm5vDpt0PM_np7cVeeRBe4-N8p7qvgsJE864gIaMXGkcWKOAQCAaAoyNEALw_wcB:G:s&s_kwcid=AL!2966!3!264955916384!!!g!437270349819!

[8] GreenSpec, ’Insulation Materials and Their Thermal Properties’, Retrieved Au-gust 29, 2019, from http://www.greenspec.co.uk/building-design/insulation-materials-thermal-properties/

[9] Home Depot, ’Owens Corning Foamular’, Retrieved August 29, 2019, from https://www.homedepot.com/p/Owens-Corning-FOAMULAR-150-1-in-x-4-ft-x-8-ft-R-5-Scored-Square-Edge-Rigid-Foam-Board-Insulation-Sheathing-20WE/207179253

[10] Interstate Plastics, ’Phenolic Sheet Natural Canvas’, Retrieved August 29, 2019, from https://www.interstateplastics.com/Phenolic-Natural-Canvas-Sheet-PHENC˜˜SH.php?sku=PHENC++SH&vid=20190923182015-2p&dim2=20&dim3=20&thickness=1.000&qty=1&recalculate.x=84&recalculate.y=1

[11] Jackson, Jelliffe. “Project Definition Document (PDD)”, University of Colorado–Boulder, Retrieved August 29,2019, from https://canvas.colorado.edu/

[12] Kaymont Balloons - Civilian / Enthusiast Balloons (HAB), from https://www.kaymont.com/habphotography

[13] Lawrence, Dale. ”Orbiting Platform for Turbulence Measurement on High Altitude Balloons for the AFOSRHYFLITS Program”, Received August 28, 2019, from https://canvas.colorado.edu/

[14] Motor Operated Valves, Retrieved August 29, 2019, from https://automationforum.co/introduction-to-motor-operated-valve-types-applications/

[15] NASA, ’State of the Art of Small Spacecraft Technology: Thermal Control’, Retrieved August 29, 2019, fromurlhttps://sst-soa.arc.nasa.gov/07-thermal

[16] Solenoid Valve, from https://www.omega.co.uk/prodinfo/solenoid-valve.html

[17] University of Colorado - Boulder Aerospace Engineering.”Past Senior Projects”, fromhttps://www.colorado.edu/aerospace/current-students/undergraduates/senior-design-projects/past-senior-projects

[18] Spincast Reel for Passive Tether Deployment, from http://www.shakespeare-fishing.com/Shakespeare-ome-casting-a-spincast-reel.html

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https://batteryuniversity.com/index.php/learn/article/discharging_at_high_and_low_temperatures

https://batteryuniversity.com/index.php/learn/article/discharging_at_high_and_low_temperatures

https://www.ncbi.nlm.nih.gov/pmc/articles/PMC5734649/

https://www.ncbi.nlm.nih.gov/pmc/articles/PMC5734649/

www.grc.nasa.gov/www/k-12/airplane/dragsphere.html

https://www.engineeringtoolbox.com/standard-atmosphere-d_604.html

https://www.engineeringtoolbox.com/standard-atmosphere-d_604.html

http://www.chem.hawaii.edu/uham/part101.html

https://www.grainger.com/product/5GCU4?gclid=Cj0KCQjww7HsBRDkARIsAARsIT45Lm5vDpt0PM_np7cVeeRBe4-N8p7qvgsJE864gIaMXGkcWKOAQCAaAoyNEALw_wcB&cm_mmc=PPC:+Google+PLA&ef_id=Cj0KCQjww7HsBRDkARIsAARsIT45Lm5vDpt0PM_np7cVeeRBe4-N8p7qvgsJE864gIaMXGkcWKOAQCAaAoyNEALw_wcB:G:s&s_kwcid=AL!2966!3!264955916384!!!g!437270349819!






http://www.greenspec.co.uk/building-design/insulation-materials-thermal-properties/

http://www.greenspec.co.uk/building-design/insulation-materials-thermal-properties/

https://www.homedepot.com/p/Owens-Corning-FOAMULAR-150-1-in-x-4-ft-x-8-ft-R-5-Scored-Square-Edge-Rigid-Foam-Board-Insulation-Sheathing-20WE/207179253



https://www.interstateplastics.com/Phenolic-Natural-Canvas-Sheet-PHENC~~SH.php?sku=PHENC++SH&vid=20190923182015-2p&dim2=20&dim3=20&thickness=1.000&qty=1&recalculate.x=84&recalculate.y=1




https://canvas.colorado.edu/

https://www.kaymont.com/habphotography

https://www.kaymont.com/habphotography

https://canvas.colorado.edu/

https://automationforum.co/introduction-to-motor-operated-valve-types-applications/

https://automationforum.co/introduction-to-motor-operated-valve-types-applications/

https://www.omega.co.uk/prodinfo/solenoid-valve.html

https://www.colorado.edu/aerospace/current-students/undergraduates/senior-design-projects/past-senior-projects

https://www.colorado.edu/aerospace/current-students/undergraduates/senior-design-projects/past-senior-projects

http://www.shakespeare-fishing.com/Shakespeare-ome-casting-a-spincast-reel.html

http://www.shakespeare-fishing.com/Shakespeare-ome-casting-a-spincast-reel.html

conceptual design document (cdd)

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