The document proposes the X-69 CargoSat space plane to deliver small satellites to low Earth orbit. It would have a crew of 2 and be able to carry 50 kg payloads. The space plane would launch from a mothership at speeds up to Mach 3.5-4.0 and deploy satellites in low Earth orbit before gliding back to land on a runway. Compared to rockets, the reusable space plane could provide more efficient and frequent delivery of small satellites without long wait times between launches.
A Presentation on Reusable Launcher Technology , with reference and basis of SpaceX Technologies Falcon 9 reusable rocket. With basic slides explaining the overview of the technology presented. ( No analytical or numerical issues addressed)
Falcon heavy Reusable Launch Vehicle- SpaceXAshish Singh
Falcon Heavy is the most powerful operational rocket in the world by a factor of two. With the ability to lift into orbit nearly 64 metric tons (141,000 lb)---a mass greater than a 737 jetliner loaded with passengers, crew, luggage and fuel--Falcon Heavy can lift more than twice the payload of the next closest operational vehicle, the Delta IV Heavy, at one-third the cost. Falcon Heavy draws upon the proven heritage and reliability of Falcon 9.
Its first stage is composed of three Falcon 9 nine-engine cores whose 27 Merlin engines together generate more than 5 million pounds of thrust at liftoff, equal to approximately eighteen 747 aircraft. Only the Saturn V moon rocket, last flown in 1973, delivered more payload to orbit. Falcon Heavy was designed from the outset to carry humans into space and restores the possibility of flying missions with crew to the Moon or Mars.
The document discusses reusable launch vehicles (RLVs) which aim to reduce the high costs of space launches by recovering and reusing rocket components. Currently, 40% of launch costs come from building non-reusable rockets. RLVs could reduce costs by a factor of 100 by recovering first stage boosters, similar to how SpaceX has landed its Falcon 9 rocket boosters. The document outlines the history of rockets, compares conventional expendable launch vehicles to reusable ones, and describes the key components and launch process of an RLV. It discusses challenges of RLVs like heat stresses during flight and challenges of vertical landing, but notes the technology is feasible and could make space travel more routine and affordable.
Falcon Heavy is SpaceX's heavy-lift launch vehicle that has the highest payload capacity of any currently operational rocket. It consists of three Falcon 9 rocket cores whose 27 Merlin engines together generate over 5 million pounds of thrust. Falcon Heavy can lift over 64 metric tons to low Earth orbit, more than any other operational vehicle, at one-third the cost of its closest competitor. It is partially reusable and designed to carry humans into space and enable missions to the Moon and Mars.
A presentation providing brief information about the Reusable Launch Vehicle or reusable rockets their past, present, and future. If you are interested in learning how this technology works do go through the slides.
This presentation briefly reviews the history of Reusable Launch Vehicle development and reuse techniques. The presentation considers a range of techniques for recovery and reuse of launch vehicles. Various different concepts of reusability have been discussed. The economics of reuse and the advantages of this technology is also presented.
This document provides an overview of the history and concepts of spaceplanes. It discusses early spaceplane designs from the 1930s and 1940s and traces the development of spaceplanes through programs like NASA's space shuttle. It describes the key characteristics of spaceplanes and the different types of spaceplane operations, including sub-orbital and single-stage to orbit. Major current and historic spaceplane designs and programs are outlined, including concepts from NASA, Airbus, Bristol Spaceplanes, Orbital Sciences, Reaction Engines, Virgin Galactic, and others. The advantages of reusable spaceplanes over expendable rockets are also summarized.
A Presentation on Reusable Launcher Technology , with reference and basis of SpaceX Technologies Falcon 9 reusable rocket. With basic slides explaining the overview of the technology presented. ( No analytical or numerical issues addressed)
Falcon heavy Reusable Launch Vehicle- SpaceXAshish Singh
Falcon Heavy is the most powerful operational rocket in the world by a factor of two. With the ability to lift into orbit nearly 64 metric tons (141,000 lb)---a mass greater than a 737 jetliner loaded with passengers, crew, luggage and fuel--Falcon Heavy can lift more than twice the payload of the next closest operational vehicle, the Delta IV Heavy, at one-third the cost. Falcon Heavy draws upon the proven heritage and reliability of Falcon 9.
Its first stage is composed of three Falcon 9 nine-engine cores whose 27 Merlin engines together generate more than 5 million pounds of thrust at liftoff, equal to approximately eighteen 747 aircraft. Only the Saturn V moon rocket, last flown in 1973, delivered more payload to orbit. Falcon Heavy was designed from the outset to carry humans into space and restores the possibility of flying missions with crew to the Moon or Mars.
The document discusses reusable launch vehicles (RLVs) which aim to reduce the high costs of space launches by recovering and reusing rocket components. Currently, 40% of launch costs come from building non-reusable rockets. RLVs could reduce costs by a factor of 100 by recovering first stage boosters, similar to how SpaceX has landed its Falcon 9 rocket boosters. The document outlines the history of rockets, compares conventional expendable launch vehicles to reusable ones, and describes the key components and launch process of an RLV. It discusses challenges of RLVs like heat stresses during flight and challenges of vertical landing, but notes the technology is feasible and could make space travel more routine and affordable.
Falcon Heavy is SpaceX's heavy-lift launch vehicle that has the highest payload capacity of any currently operational rocket. It consists of three Falcon 9 rocket cores whose 27 Merlin engines together generate over 5 million pounds of thrust. Falcon Heavy can lift over 64 metric tons to low Earth orbit, more than any other operational vehicle, at one-third the cost of its closest competitor. It is partially reusable and designed to carry humans into space and enable missions to the Moon and Mars.
A presentation providing brief information about the Reusable Launch Vehicle or reusable rockets their past, present, and future. If you are interested in learning how this technology works do go through the slides.
This presentation briefly reviews the history of Reusable Launch Vehicle development and reuse techniques. The presentation considers a range of techniques for recovery and reuse of launch vehicles. Various different concepts of reusability have been discussed. The economics of reuse and the advantages of this technology is also presented.
This document provides an overview of the history and concepts of spaceplanes. It discusses early spaceplane designs from the 1930s and 1940s and traces the development of spaceplanes through programs like NASA's space shuttle. It describes the key characteristics of spaceplanes and the different types of spaceplane operations, including sub-orbital and single-stage to orbit. Major current and historic spaceplane designs and programs are outlined, including concepts from NASA, Airbus, Bristol Spaceplanes, Orbital Sciences, Reaction Engines, Virgin Galactic, and others. The advantages of reusable spaceplanes over expendable rockets are also summarized.
This document discusses reusable launch vehicles (RLVs). It begins with an introduction that defines RLVs as vehicles that can be used for multiple missions. The main advantage is lower costs compared to expendable rockets. The history section discusses early concepts from the 1950s and serious attempts in the 1990s by companies like McDonnell-Douglas and Lockheed Martin. The present section notes SpaceX's success in recovering Falcon 9 first stages. Design considerations for RLVs include withstanding high stresses and temperatures during launch and reentry. Stages to orbit discusses single-stage and multi-stage options. Vertical landing and retro-propulsion methods are also covered. Preparing a reused RLV requires extensive inspection and refurbishment of components
SpaceX’s Falcon 9 and Blue Origin Reusable Launch Vehicles are designed not only to withstand re-entry but also to return to the launch pad or ocean landing site for a vertical landing. Reusable rocket is the pivotal breakthrough needed to substantially reduce the cost of space access and make human multi-planet species
ISRO is developing a reusable launch vehicle technology demonstrator (RLV-TD) to test technologies for reusable two-stage orbital launch systems. The RLV-TD will be launched on a solid booster and glide back to land on a runway, testing hypersonic flight and autonomous landing capabilities. Wind tunnel testing of designs has been completed. The first flight experiment (HEX) will involve ocean recovery of telemetry data, followed by additional tests to validate powered cruise flight and horizontal runway landing (LEX), and eventually orbital demonstrations. The RLV-TD launch is planned for 2014 and aims to significantly reduce the cost of launching payloads to space.
The document discusses India's GSLV Mark-III heavy lift launch vehicle. It has three stages - two solid rocket boosters as the first stage, a liquid propellant core as the second stage, and a cryogenic upper stage as the third stage. The document outlines the key features and thrust of each stage. It notes that the successful test of the CARE module, which separated from the cryogenic stage and landed safely, verifies technologies for ISRO's human spaceflight program to send astronauts into low Earth orbit. The first orbital flight of GSLV Mark-III is planned for 2016.
This document discusses various concepts for propulsion systems to enable future space missions. It covers existing expendable launch vehicles, shuttle-derived vehicles using components of the space shuttle, and more advanced concepts like nuclear thermal rockets, solar sails, and fusion propulsion. Key shuttle-derived concepts mentioned include Shuttle-C, which would replace the orbiter with a cargo canister, and Shuttle-B, which would use expendable engines attached to the external tank. The document also discusses in-space propulsion options, including Project Prometheus to develop nuclear-electric propulsion.
The document discusses India's GSLV-MK III launch vehicle and its upcoming Crew Module Atmospheric Re-entry Experiment (CARE). The GSLV-MK III will be capable of launching satellites into geostationary orbit and is intended to launch an Indian crew vehicle. On December 18, 2014, ISRO will conduct an experimental suborbital launch of the GSLV MK-3 without its cryogenic third stage to test the Crew Module through atmospheric re-entry and splashdown in the Bay of Bengal.
The document discusses India's GSLV Mark III rocket. Key points:
- GSLV Mark III is India's most powerful rocket to date and an important step for sending Indian astronauts to space.
- On a test flight, it will carry a crew module prototype and validate the rocket's performance during atmospheric ascent.
- The rocket can carry over 10 metric tons to low Earth orbit and over 4 metric tons to geostationary transfer orbit.
- The test flight will check the performance of the rocket's two powerful solid rocket boosters and twin-engine liquid fuel first stage.
Presentation by Steve Cook at the AAS Von Braun Memorial Symposium in Huntsville, Alabama, 21 October 2008.
<a href="http://astronautical.org/vonbraun/vonbraun-2008/session1">http://astronautical.org/vonbraun/vonbraun-2008/session1</a>
The document summarizes information about satellite launch vehicles, including their origin from ballistic missiles, types (expendable vs reusable), how they work based on Newton's 3rd law of motion, common fuels, ideal launch bases near the equator, and details about India's satellite launch vehicles - SLV, ASLV, and PSLV. It provides an overview of key concepts about satellite launch vehicles.
The document summarizes an upcoming Atlas V 401 launch that will deliver a GPS IIF-12 satellite to semi-synchronous orbit. It will launch from Cape Canaveral Air Force Station, Florida. The GPS satellite is part of a constellation that provides navigation data to users worldwide. The Atlas V 401 rocket has successfully launched many missions, including previous GPS satellites.
Presentation by Clinton Dorris (Deputy Manager, Altair Project Office, NASA) at the Von Braun Memorial Symposium in Huntsville, Alabama, 21 October 2008.
<a href="http://astronautical.org/vonbraun/vonbraun-2008/session2">http://astronautical.org/vonbraun/vonbraun-2008/session2</a>
The document provides details about SpaceX's Falcon 9 Reusable (F9R) launch vehicle. It describes the key aspects of the F9R including its first and second stages, 9-engine configuration, reusable first stage capabilities using landing legs and grid fins, payload capabilities, and upcoming missions. The goal of the F9R is to revolutionize space access by providing a fully and rapidly reusable rocket to substantially reduce launch costs.
The document proposes an ultralight solar-powered hybrid research drone for exploring Mars. The drone would use lighter-than-air gases like hydrogen or helium inside its lenticular hull to achieve buoyancy and reduce structural mass. It can take off and land vertically and transition to efficient horizontal cruise flight using its wings, engines, and moveable control surfaces. The drone is designed to operate autonomously using solar energy from photovoltaic cells on its hull to recharge batteries and potential fuel cells. It combines aspects of airplanes, airships, and helicopters into a single vehicle capable of performing scientific observations across the Martian surface with minimal infrastructure requirements.
India began developing launch vehicles in the 1970s, starting with the SLV-3 experimental satellite launch vehicle. The Augmented Satellite Launch Vehicle (ASLV) was successfully launched in 1992. Key vehicles developed include the Polar Satellite Launch Vehicle (PSLV) and Geosynchronous Satellite Launch Vehicle (GSLV). The PSLV has had 24 consecutive successful flights out of 25 launches and can launch satellites into sun-synchronous and geo-synchronous orbits. The GSLV can launch 2-2.5 tonne satellites into geo-synchronous transfer orbit, with 4 successful flights out of 7. India is also developing the larger GSLV Mark III to enable launching of heavier 4.5-5 tonne satellites.
Paper - SmallSat 1999 Management ChallengesDave Callen
This document discusses the management challenges of launching multiple payloads for multiple customers using Orbital Sciences Corporation's Taurus launch vehicle. It describes Orbital's approach of providing integrated launch services to multiple customers on a single launch in order to reduce costs. This adds complexities of satisfying different customer requirements, schedules, and technical and political issues. The document outlines some of the challenges Orbital has faced in managing shared launches and coordinating between different customer payloads. It also provides examples of previous Orbital missions with multiple payloads launched using different configurations on the Pegasus and Taurus vehicles.
Information regarding SPACE TRANSPORTATION SYSTEM {SPACE SHUTTLE} and its Robotic arm , THERMAL PROTECTION SYSTEM With upcoming ORION(MPCV).
POINT OF TALK:
1.Introduction to space shuttle.
2.Description
3. photo with parts and TILES(TPS)
4. Various stages from takeoff to landing
5.COLLECTING OF EXTERNAL TANK
6.ROBOTIC ARM OF SPACE SHUTTLE
7.SPACE SHUTTLE DURING RE-ENTRY
8.THERMAL PROTECTION SYSTEM
9.REUSABLE CERAMIC TILES
10.COMPOSITION OF TILES
11.TESTING OF TILES
12..ORION MPCV
Presentation by Phil Sumrall (Advanced Planning Manager, NASA) at the Von Braun Memorial Symposium in Huntsville, Alabama, 21 October 2008.
<a href="http://astronautical.org/vonbraun/vonbraun-2008/session2">http://astronautical.org/vonbraun/vonbraun-2008/session2</a>
The document discusses the Geosynchronous Satellite Launch Vehicle Mark III (GSLV Mk III) developed by the Indian Space Research Organization (ISRO). It has a higher payload capacity than the GSLV Mk II and can carry 4 ton class satellites to Geosynchronous Transfer Orbit or 10 tons to Low Earth Orbit. The vehicle uses two solid strap-on boosters and a liquid core booster with a cryogenic upper stage. GSLV Mk III has successfully launched satellites like GSAT-19, GSAT-29 and the Chandrayaan-2 lunar mission.
LA PROFECIA Y PROMESA DE DIOS HECHA AL PUEBLO DE ISRAEL, EN BOCA DEL PROFETA ISAIAS , SE CUMPLE.
ASI LOCITA MATEO, TODO ESTO PASO PARA QUE SE CUMPLIERA LO DICHO POR EL PROFETA.
This document discusses reusable launch vehicles (RLVs). It begins with an introduction that defines RLVs as vehicles that can be used for multiple missions. The main advantage is lower costs compared to expendable rockets. The history section discusses early concepts from the 1950s and serious attempts in the 1990s by companies like McDonnell-Douglas and Lockheed Martin. The present section notes SpaceX's success in recovering Falcon 9 first stages. Design considerations for RLVs include withstanding high stresses and temperatures during launch and reentry. Stages to orbit discusses single-stage and multi-stage options. Vertical landing and retro-propulsion methods are also covered. Preparing a reused RLV requires extensive inspection and refurbishment of components
SpaceX’s Falcon 9 and Blue Origin Reusable Launch Vehicles are designed not only to withstand re-entry but also to return to the launch pad or ocean landing site for a vertical landing. Reusable rocket is the pivotal breakthrough needed to substantially reduce the cost of space access and make human multi-planet species
ISRO is developing a reusable launch vehicle technology demonstrator (RLV-TD) to test technologies for reusable two-stage orbital launch systems. The RLV-TD will be launched on a solid booster and glide back to land on a runway, testing hypersonic flight and autonomous landing capabilities. Wind tunnel testing of designs has been completed. The first flight experiment (HEX) will involve ocean recovery of telemetry data, followed by additional tests to validate powered cruise flight and horizontal runway landing (LEX), and eventually orbital demonstrations. The RLV-TD launch is planned for 2014 and aims to significantly reduce the cost of launching payloads to space.
The document discusses India's GSLV Mark-III heavy lift launch vehicle. It has three stages - two solid rocket boosters as the first stage, a liquid propellant core as the second stage, and a cryogenic upper stage as the third stage. The document outlines the key features and thrust of each stage. It notes that the successful test of the CARE module, which separated from the cryogenic stage and landed safely, verifies technologies for ISRO's human spaceflight program to send astronauts into low Earth orbit. The first orbital flight of GSLV Mark-III is planned for 2016.
This document discusses various concepts for propulsion systems to enable future space missions. It covers existing expendable launch vehicles, shuttle-derived vehicles using components of the space shuttle, and more advanced concepts like nuclear thermal rockets, solar sails, and fusion propulsion. Key shuttle-derived concepts mentioned include Shuttle-C, which would replace the orbiter with a cargo canister, and Shuttle-B, which would use expendable engines attached to the external tank. The document also discusses in-space propulsion options, including Project Prometheus to develop nuclear-electric propulsion.
The document discusses India's GSLV-MK III launch vehicle and its upcoming Crew Module Atmospheric Re-entry Experiment (CARE). The GSLV-MK III will be capable of launching satellites into geostationary orbit and is intended to launch an Indian crew vehicle. On December 18, 2014, ISRO will conduct an experimental suborbital launch of the GSLV MK-3 without its cryogenic third stage to test the Crew Module through atmospheric re-entry and splashdown in the Bay of Bengal.
The document discusses India's GSLV Mark III rocket. Key points:
- GSLV Mark III is India's most powerful rocket to date and an important step for sending Indian astronauts to space.
- On a test flight, it will carry a crew module prototype and validate the rocket's performance during atmospheric ascent.
- The rocket can carry over 10 metric tons to low Earth orbit and over 4 metric tons to geostationary transfer orbit.
- The test flight will check the performance of the rocket's two powerful solid rocket boosters and twin-engine liquid fuel first stage.
Presentation by Steve Cook at the AAS Von Braun Memorial Symposium in Huntsville, Alabama, 21 October 2008.
<a href="http://astronautical.org/vonbraun/vonbraun-2008/session1">http://astronautical.org/vonbraun/vonbraun-2008/session1</a>
The document summarizes information about satellite launch vehicles, including their origin from ballistic missiles, types (expendable vs reusable), how they work based on Newton's 3rd law of motion, common fuels, ideal launch bases near the equator, and details about India's satellite launch vehicles - SLV, ASLV, and PSLV. It provides an overview of key concepts about satellite launch vehicles.
The document summarizes an upcoming Atlas V 401 launch that will deliver a GPS IIF-12 satellite to semi-synchronous orbit. It will launch from Cape Canaveral Air Force Station, Florida. The GPS satellite is part of a constellation that provides navigation data to users worldwide. The Atlas V 401 rocket has successfully launched many missions, including previous GPS satellites.
Presentation by Clinton Dorris (Deputy Manager, Altair Project Office, NASA) at the Von Braun Memorial Symposium in Huntsville, Alabama, 21 October 2008.
<a href="http://astronautical.org/vonbraun/vonbraun-2008/session2">http://astronautical.org/vonbraun/vonbraun-2008/session2</a>
The document provides details about SpaceX's Falcon 9 Reusable (F9R) launch vehicle. It describes the key aspects of the F9R including its first and second stages, 9-engine configuration, reusable first stage capabilities using landing legs and grid fins, payload capabilities, and upcoming missions. The goal of the F9R is to revolutionize space access by providing a fully and rapidly reusable rocket to substantially reduce launch costs.
The document proposes an ultralight solar-powered hybrid research drone for exploring Mars. The drone would use lighter-than-air gases like hydrogen or helium inside its lenticular hull to achieve buoyancy and reduce structural mass. It can take off and land vertically and transition to efficient horizontal cruise flight using its wings, engines, and moveable control surfaces. The drone is designed to operate autonomously using solar energy from photovoltaic cells on its hull to recharge batteries and potential fuel cells. It combines aspects of airplanes, airships, and helicopters into a single vehicle capable of performing scientific observations across the Martian surface with minimal infrastructure requirements.
India began developing launch vehicles in the 1970s, starting with the SLV-3 experimental satellite launch vehicle. The Augmented Satellite Launch Vehicle (ASLV) was successfully launched in 1992. Key vehicles developed include the Polar Satellite Launch Vehicle (PSLV) and Geosynchronous Satellite Launch Vehicle (GSLV). The PSLV has had 24 consecutive successful flights out of 25 launches and can launch satellites into sun-synchronous and geo-synchronous orbits. The GSLV can launch 2-2.5 tonne satellites into geo-synchronous transfer orbit, with 4 successful flights out of 7. India is also developing the larger GSLV Mark III to enable launching of heavier 4.5-5 tonne satellites.
Paper - SmallSat 1999 Management ChallengesDave Callen
This document discusses the management challenges of launching multiple payloads for multiple customers using Orbital Sciences Corporation's Taurus launch vehicle. It describes Orbital's approach of providing integrated launch services to multiple customers on a single launch in order to reduce costs. This adds complexities of satisfying different customer requirements, schedules, and technical and political issues. The document outlines some of the challenges Orbital has faced in managing shared launches and coordinating between different customer payloads. It also provides examples of previous Orbital missions with multiple payloads launched using different configurations on the Pegasus and Taurus vehicles.
Information regarding SPACE TRANSPORTATION SYSTEM {SPACE SHUTTLE} and its Robotic arm , THERMAL PROTECTION SYSTEM With upcoming ORION(MPCV).
POINT OF TALK:
1.Introduction to space shuttle.
2.Description
3. photo with parts and TILES(TPS)
4. Various stages from takeoff to landing
5.COLLECTING OF EXTERNAL TANK
6.ROBOTIC ARM OF SPACE SHUTTLE
7.SPACE SHUTTLE DURING RE-ENTRY
8.THERMAL PROTECTION SYSTEM
9.REUSABLE CERAMIC TILES
10.COMPOSITION OF TILES
11.TESTING OF TILES
12..ORION MPCV
Presentation by Phil Sumrall (Advanced Planning Manager, NASA) at the Von Braun Memorial Symposium in Huntsville, Alabama, 21 October 2008.
<a href="http://astronautical.org/vonbraun/vonbraun-2008/session2">http://astronautical.org/vonbraun/vonbraun-2008/session2</a>
The document discusses the Geosynchronous Satellite Launch Vehicle Mark III (GSLV Mk III) developed by the Indian Space Research Organization (ISRO). It has a higher payload capacity than the GSLV Mk II and can carry 4 ton class satellites to Geosynchronous Transfer Orbit or 10 tons to Low Earth Orbit. The vehicle uses two solid strap-on boosters and a liquid core booster with a cryogenic upper stage. GSLV Mk III has successfully launched satellites like GSAT-19, GSAT-29 and the Chandrayaan-2 lunar mission.
LA PROFECIA Y PROMESA DE DIOS HECHA AL PUEBLO DE ISRAEL, EN BOCA DEL PROFETA ISAIAS , SE CUMPLE.
ASI LOCITA MATEO, TODO ESTO PASO PARA QUE SE CUMPLIERA LO DICHO POR EL PROFETA.
O projeto tem como objetivo principal reunir pessoas interessadas em debater arte e compartilhar estas reflexões, com os objetivos secundários de criar um espaço de intercâmbio entre iniciantes e experientes sobre arte, e tornar a universidade uma referência local nesta discussão. Será realizado em encontros quinzenais para debates e apresentações sobre obras de arte, com a fase de divulgação incluindo pôsteres, cartões postais, palestras e publicações para difundir as análises produzidas.
Este documento presenta los requisitos para un trabajo de investigación individual sobre blogs. Los estudiantes deben incluir definiciones de blogs de 3 fuentes, las diferencias entre blogs, páginas web y foros, las características y elementos de los blogs, programas gratuitos para crear blogs, y curiosidades sobre blogs. El trabajo debe entregarse en PowerPoint antes del 26 de noviembre incluyendo portada, introducción, conclusión, índice y bibliografía.
Placas de advertência têm por finalidade alertar os usuários da via sobre condições potencialmente perigosas, indicando sua natureza, e suas mensagens possuem caráter de recomendação.
Dora Reyes Mani is the subject of this document from the Instituto Eulogio Gillow dated February 2012. The document appears to be related to Dora Reyes Mani and the Instituto Eulogio Gillow from February 2012 but provides no other details in the short text.
El documento presenta un proyecto de una alumna de cuarto año sobre isometría y proyección ortogonal de figuras. La alumna Massiel Caballero de la sección D realiza un estudio sobre isometría y proyección ortogonal de objetos y proporciona un enlace a un video de YouTube como referencia adicional.
We are living in a vast universe that contains tremendous unknown knowledge. Human space exploration helps to address the fundamental questions about our place in the universe. In this the development of spacecrafts is remarkable. SKYLON is space plane that can be a replacement for the current scenario of space travel by its reliability, ease of operation and economic friendly nature. It’s a single stage to orbit hypersonic space plane. That uses horizontal take off and landing like a conventional aircraft. It could reach up to the low earth orbit (LEO) with a payload of about 15 tons. This system use combined cycle engine commonly known as synergistic air breathing rocket engine (SABRE).That works both in air breathing and pure rocket mode. This permits the vehicle to cruise at hypersonic speed (around Mach 5.5) within earth atmosphere. SKYLON is the future of aviation and space industry, which may ease many missions from earth surface to space. Further modification in the engine may lead not only to the orbit but also far away from that .its low fuel consumption lower weight and reduced risk factor increases the performance and makes possible space tourism for people belongs to any community
Propulsion System in Hypersonic Spacecraft Rocket: A Review of Recent Develop...IRJET Journal
This document provides a review of propulsion systems for hypersonic spacecraft. It discusses the history of hypersonic vehicle development from early concepts in the 1930s to current programs. Various propulsion technologies are described, including nuclear thermal rockets, air-breathing engines like scramjets, and combinations of ducted jet and rocket engines. Challenges in developing efficient hypersonic propulsion systems are also reviewed.
This document provides information about reusable launch vehicles, specifically SpaceX's Falcon 9 Reusable (F9R) launch vehicle. It discusses SpaceX's goals of making spaceflight routine and affordable. The document describes the Falcon 9 family of launch vehicles and provides details on the technical specifications and design of the Falcon 9R first and second stages. It also discusses the payload fairing and Dragon spacecraft. Applications and advantages of reusable launch vehicles are presented.
This document summarizes a trade study conducted to determine the best design for a deployable drag device to accelerate the orbital decay of upper stage launch vehicles. The trade study compared different device types (drag sails, inflatables, tethers, propulsive), launch vehicles, target orbit altitudes, desired decay times, and packaging constraints. A drag sail design was proposed as the baseline option due to its lower cost and mass efficiency compared to a propulsive option, while still meeting the design requirement of deorbiting within 25 years. A stability analysis was then conducted to determine the optimal drag sail configuration before preliminary component selection and structural analysis.
This document describes a proposed horizontal in-line launch staging (HILLS) system. HILLS involves joining two delta wing aircraft nose to tail to operate as a single reusable vehicle, with the first stage powering both until separation to deliver the second stage and payloads to orbit. Combined with other technologies, HILLS could allow affordable reusable launch services with lessons from the Concorde and lifting bodies providing economy, safety, low vibration and low g-forces. The company is developing prototypes like the SKIP airframe to demonstrate technologies for eventual orbital vehicles.
Do you want to go into space? To float weightless? To look down at the whole Earth from the blackness of space? It’s an exciting time for private space travel. Many new space companies have been founded in just the last few years. Most are working toward the goal of ordinary people going to space.
This presentation summarizes current space tourism programs, from weightless rides in a Zero G plane to trips to the International Space Station. It also presents an overview of near-future private space projects, including SpaceShipTwo suborbital rocket flights, orbital rides in the Dragon capsule, and the private Genesis space hotel currently being built in orbit.
RLV.pptx and charging system in electrical vehicleSuruAarvi
The document discusses the technology of reusable launch vehicles for satellites. It begins with an introduction to reusable launch vehicles and their main components. It then discusses the history of reusable launch vehicle development. Current programs from SpaceX, ISRO, and other organizations are presented. The working mechanisms from launch to landing of reusable launch vehicles is explained. Key aspects of reusable launch vehicle design like stages to orbit, vertical landing, and retro-propulsion are outlined. Finally, the economics of reuse and technologies required for reusable launch vehicle feasibility are presented.
Project LEON proposes utilizing ram accelerator technology to cheaply launch small satellites into low Earth orbit. The ram accelerator works by accelerating projectiles through long tubes using combustion, reaching speeds to deploy payloads into orbit within minutes. This provides a lower-cost alternative to rockets for launching small satellites. Project LEON aims to enable more frequent and affordable launches for purposes like communications networks and removing space debris. The ram accelerator technology has been demonstrated at universities and could lower launch costs by an order of magnitude compared to rockets.
This document discusses reusable launch vehicles (RLVs) and their potential to significantly reduce the cost of space travel and access. It provides details on some of the major players in RLV development, including SpaceX, Blue Origin, and ISRO. SpaceX's Falcon 9 is highlighted as an example of a partially reusable orbital launch system. The document argues that fully reusable single-stage-to-orbit rockets, if achievable, could reduce launch costs by as much as 100 times and revolutionize space exploration and satellite deployment.
Skylon spaceplane, uk the spacecraft of tomorrowhindujudaic
Skylon is a design for a single-stage-to-orbit spaceplane by Reaction Engines Limited that uses a novel SABRE engine. The SABRE engine allows Skylon to accelerate to Mach 5.4 using atmospheric oxygen before switching to onboard liquid oxygen for orbital insertion. Skylon could carry 15 tons to low Earth orbit on each trip and return to Earth for reuse within two days. If successful, Skylon could begin test flights in 2019 and visit the International Space Station by 2022, dramatically reducing the cost of launching cargo to orbit.
The document discusses various new space companies and concepts in commercial spaceflight. It describes NewSpace as a term for new space companies that take different approaches to space development than traditional aerospace companies, focusing on private funding, low-cost approaches, and innovation. It then profiles over 15 NewSpace companies, describing their goals and projects, which include suborbital tourism ventures, rocket development programs, and work on reusable launch vehicles and spacecraft. It also discusses several space competitions and organizations working in this emerging commercial space industry.
Presentation by Lawrence Williams (SpaceX) at the Von Braun Memorial Symposium in Huntsville, Alabama, 22 October 2008.
<a href="http://astronautical.org/vonbraun/vonbraun-2008/session5">http://astronautical.org/vonbraun/vonbraun-2008/session5</a>
This document discusses the potential for using airships as science platforms for Earth and space science. It proposes the 20-20-20 Airships Challenge through NASA's Centennial Challenges program to encourage development of stratospheric airships. The challenge would award prizes for airships that can carry payloads of 20kg to 20km altitude for 20 hours (Tier 1) or 200kg to 20km altitude for 200 hours (Tier 2). Airships could enable new types of long duration observations for Earth science, atmospheric science, and astrophysics at lower costs than current space missions. The document reviews some example science instruments and investigations that could be performed from airship platforms.
The document summarizes the Hypersonic Airplane Space Tether Orbital Launch (HASTOL) concept studied by Boeing and NASA. It describes the Phase I and Phase II studies, which evaluated the feasibility of using a hypersonic aircraft launched from a mothership to rendezvous with a tether in high altitude flight and transfer a payload to low Earth orbit. Key findings from Phase I included validation that a tether tip could withstand thermal loads from dipping into the atmosphere and identification of a rotovator tether concept. Phase II involved further development of system requirements, conceptual design including of an air-launched turbo-rocket hypersonic plane, and analysis of potential mission opportunities and markets.
LauncherOne revolutionary orbital transport for small satellitesDmitry Tseitlin
Virgin Galactic offers a new small satellite launch vehicle called LauncherOne that will provide an affordable, dedicated, and responsive ride to orbit for small satellites. LauncherOne will be air-launched from Virgin Galactic's carrier aircraft WhiteKnightTwo and will be capable of delivering a 500 lb payload to low Earth orbit, addressing the growing needs of the microsatellite market. By leveraging the technologies and infrastructure developed for its SpaceShipTwo suborbital space plane, Virgin Galactic aims to launch LauncherOne in 2016 and provide a low-cost option for small satellite operators to access space.
AM PRESENTING U MA SEMINAR SLIDES ON TOPIC "HYPERSONIC AIR BREATHING ENGINES" ROOTED UP BY HELP OF NASA INFORMATION.SINCE I AM INTERESTED IN SPACE STUDIES I CHOOSE THIS,EVENTHOUGH AM A MECH ENGINEER!! ..I KNOW , SOMEONE OR ANYONE BE GAINFUL BY THIS.....DURING MA SEMINAR I HOLD ON MANY SITES TO PROVIDE RELATD SLIDES,BUT THEY ALL NEED REGISTRATION,MONEY AND ALL...BUT ITZ NOT FAIR.!!...SO AM SHARING U WITH THIS.....FOR ANY DOUBTS OR REPORTS,SUPPORTING JOURNELS ,CONTACT ME: sanoojsiddikh@gmail.com
SpinLaunch is developing a new kinetic launch system that uses a large centrifuge to accelerate payloads to hypersonic speeds before launch, reducing the need for chemical rockets by 70%. The system works by spinning a payload up to 5,000 mph inside a large vacuum chamber and then releasing it. A small rocket then provides the final boost to orbit. SpinLaunch successfully tested a 1/3 scale prototype in 2021, launching a projectile tens of thousands of feet into the atmosphere. If successful at full scale, the system could reduce launch costs by 20-fold to $500,000 per launch and allow for more frequent daily launches compared to conventional rockets. However, payloads will be limited to around 440 pounds and the g
1) The document discusses recent progress and developments in 2016 related to hypersonic flight technologies, including experimental hypersonic vehicles and engines.
2) Key programs discussed include the HAWC hypersonic missile program, Lockheed Martin's SR-72 spy plane concept, and Europe's LAPCAT program developing hypersonic transports.
3) Advances are being made in scramjet and air-breathing propulsion technologies to enable aircraft and launch vehicles capable of hypersonic flight between Mach 5-8 speeds for applications in reconnaissance, weapons, and potential future civil transports crossing continents within hours.
This document summarizes Robert White's qualifications for an aerospace engineering position. It includes his education background with a B.S. in Aerospace Engineering and minor in Electrical Engineering from Purdue University. It outlines relevant skills and work experience, including internships modifying balloon and robot camera systems. It also details involvement in student engineering projects including a bucket elevator excavator for a lunar mining robot and cargo lander designs for a moon base project.
Heat Transfer Analysis for a Winged Reentry Flight Test BedCSCJournals
In this paper we deal with the aero-heating analysis of a reentry flight demonstrator helpful to the research activities for the design and development of a possible winged Reusable Launch Vehicle. In fact, to reduce risks in the development of next generation reusable launch vehicles, as first step it is suitable to gain deep design knowledge by means of extensive numerical computations, in particular for the aero-thermal environment the vehicle has to withstand during reentry. The demonstrator under study is a reentry space glider, to be used both as Crew Rescue Vehicle and Crew Transfer Vehicle for the International Space Station. It is designed to have large atmospheric manoeuvring capability, to test the whole path from the orbit down to subsonic speeds and then to the landing on a conventional runway. Several analysis tools are integrated in the framework of the vehicle aerothermal design. Between the others, we used computational analyses to simulate aerothermodynamic flowfield around the spacecraft and heat flux distributions over the vehicle surfaces for the assessment of the vehicle Thermal Protection System design. Heat flux distributions, provided for equilibrium conditions of radiation at wall and thermal shield emissivity equal to 0.85, highlight that the vehicle thermal shield has to withstand with about 1500 [kW/m2] and 400 [kW/m2] at nose and wing leading edge, respectively. Therefore, the fast developing new generation of thermal protection materials, such as Ultra High Temperature Ceramics, are available candidate to built the thermal shield in the most solicited vehicle parts. On the other hand, away from spacecraft leading edges, due to the low angle of attack profile followed by the vehicle during descent, the heat flux is close to values attainable with conventional heat shield. Also, the paper shows that the flying test bed is able to validate hypersonic aerothermodynamic design database and passenger experiments, including thermal shield and hot structures, giving confidence that a full-scale development can successfully proceed.
Heat Transfer Analysis for a Winged Reentry Flight Test Bed
X-69
1. X-69 CargoSat Space-Plane for LEO deliveries
Introduction:
It is quite often to use cryogenic rockets to deliver various satellites into space. Due to size
constraints of rockets, their payload, propulsive efficiencies, they house all the satellites together with a
strategy to launch them at once which keeps customers waiting till all the tickets are booked even if some
of them happen to be small sized satellites. This project is an attempt to use X-69 CargoSat space-plane
instead of rockets to deliver those small satellites individually or together if accommodated without long
waiting and required efficiencies.
2.1 Mission Specifications:
Table 1: Anticipated Specifications
General Characteristics
Crew 2
Payload 110.23 lbs (50 kg)
Loaded weight 10,000 kg (22046.23 lbs)
Powerplant 1x RocketmotorTwo liquid/solid hybrid rocket engine
Performance
Maximum speed 4,000 km/hr (2,500 mph)
Orbit Low Earth Orbit
Mach 3.5 – 4.0
Launch and Landing Characteristics
Launch Vehicles B-52 Stratofortress, White Knight Two
Launch Speed Approx. 0.7 – 0.8
Launch Altitude/Service Ceiling 70,000 ft (21,000 m)
Landing distance
2. 2.2 Mission Profile:
Mission profile for X-69 will look similar to that of X-15 as shown in Fig. 1. There will be
modifications after X-69 dives into the space, although its in-atmosphere is similar to that of X-15. Any
mothership will help to air launch X-69. After detachment, it will burnout and climb towards LEO using
RocketMotorTwo engine or similar ones.
Fig 2. shows anticipated mission profile of X-69 CargoSat. Depending on mission it may spend
variable amount of time in space either to just deploy and or wait for docking back another satellite. Re-
entry will initiated as it drops into the atmosphere and will glide and land on airport.
Fig. 1. Mission Profile of X-15
3. Fig 2. Estimated Mission profile of X-69 CargoSat
2.3 Market Analysis:
Since a decade, concept of cubesats and other small satellites has been trending with a rising market.
They deliver crucial advantages like compactness, multifunctional characteristics due to advanced
technologies and highly efficient features. To date, rockets have been used to deliver these satellites into
the space or orbits. Rockets are good to transport heavy and large instruments into the space. Although if a
company or a customer wants to put their small-sized satellites to space, they have to wait or pay more price
if they seek to launch through rockets. Moreover, it keeps them waiting till all other satellite companies
collaborate for launch.
4. Reusable space-planes have capabilities to reach up to LEO and further. Small satellites can be
delivered into space by space-planes with reusable capabilities. X-planes like Virgin Galactic’s Spaceship
Two, X-15, X-37B can make return trips with crew inside.
2.4 Technical and Economic Feasibility
Space-planes carrying human to space have been under research and even been flown quite many
times. From technical perspective, it would be feasible to re-design the space-planes for satellite transport
with reusability which essentially gives the benefit for aborted missions. These planes can also be used to
bring back damaged or reparable satellites with efficient re-entry. Following points can be considered to
propose the design of X-69 CargoSat:
Based on payload and other avionics features, the design of X-69 will be similar to that been
trending for Space-planes for human space flight.
Air-launch will be same, using either B-52 Stratofortress or White Knight Two aircrafts.
Will have to re-consider the aerodynamics for re-entry and landing aspects.
Propulsion system will change based on payload, carrying satellites back and forth.
Reusability is always an advantage for space transportation. Moreover, many attempts have been
made and some of them are even successful to bring back rockets from space like SpaceX and Blue Origin.
Although, this takes lot of fuel to fight with re-entry speeds and again to manage a perfect landing. It is
quite easy for an airplane-like structure to land with less maintenance and accuracy. Again, space-planes
would not need specific launch/landing pad. If X-69 brings reparable satellites from space, it can be
delivered wherever it is necessary with less earth-transportation issues. It can land on any airport, deliver
the satellite.
For instance, let’s say a Chinese Space agency collaborates and asks American space agency who
makes X-69 to bring back their damaged satellite, it will be easy and feasible to rendezvous X-69 to that
satellite, dock it in, re-enter and land on any airport in China delivering the payload and flying back to home
country. This will reduce ground transport cost, will not have to about damaged satellites and many other
advantages.
2.5 Critical Mission Requirements:
Following are the critical mission requirements that should be considered for design:
Delta-wing design for re-entry and efficient landing.
Efficient propulsion system for X-69 while air launching.
Outer body material to deal with re-entry heat and high temperatures.
Delta-wing pattern is quite traditional for supersonic aircrafts. Its design depends on requirements
of aircraft and other technical specifications like altitude, speed, take-off and landing distance etc. Due to
hypersonic speeds at re-entry, the surround atmospheric air heats up which is unfavorable for aircrafts with
normal body material. Using carbon-composite makes it light-weight, resistant to high temperature and
pressure and many other structural advantages.
5. 3.0 Comparative Study of Similar Airplanes:
Table 2: Comparative study
Parameters X-69 CargoSat Virgin Galactic’s
Spaceship Two
Boeing’s X-37 Boeing’s X-20
Dyna-Soar
X-15
Crew 2 2 crew and 6
passengers
none 1 pilot 1 pilot
Takeoff/Launch
weight
22,000 lb 21,428 lb (9,740 kg) 11,000 lb
(4,990 kg)
11,387 lb (5,165
kg)
34,000 lb (15,420 kg)
Empty weight 15,000 lb (6,804 kg) 15,000 lb (6,804 kg) NA (electric
powered)
10,395 lb (4,715
kg)
14,600 lb (6,620 kg)
Thrust 60,000 lbf to 75,000
lbf
60,000 lbf (270 kN) 157.4 lbf 700
kN)
72,000 lbf (323
kN)
70,400 lbf (313 kN)
Critical Speed, Vcr 2,500 mph 2,500 mph (4,000
km/hr)
(Orbital)
17,426 mph
(28,440 km/h)
17,500 mph
(28,165 km/hr)
4,520 mph (7,274 km/h)
Range, R 300 mi and apogee of
upto LEO of 160 –
250 km
Planned apogee of
110 km
Earth orbit 22,000
nm (40,700 km)
280 mi (450 km)
Wing Area, S 345 ft2
(32 m2
) 200 ft2
(18.6 m2
)
Wing span, b 27 ft (8.3 m) 14 ft 11 in (4.5
m)
20 ft 10 in (6.34
m)
22 ft 4 in (6.8 m)
Aspect Ratio, AR 1.256 2.486
Type of Payload Crew and satellite Crew Satellites Crew Crew
Powerplant 1x Rocket motorTwo
liquid/solid hybrid
rocket engine
Gallium
Arsenide Solar
Cells with Li-
Ion batteries
1x Transtage
rocket engine
1x Reaction Motors
XLR99-RM-2 liquid
propellant rocket engine
3.3 Discussion:
Various design The design of X-69 will be challenging from various aerospace aspects. There is
quite a lot of versatility in comparing X-69 with other Space-planes considering their performance, general
characteristics, applications, etc. As discussed, its launch weight will be similar to that of SpaceshipTwo of
about 22,000 lb. From the comparison table, optimal thrust will be considered in the range of 60,000 to
75,000 lbf based on payload. X-15 using XLR99-RM-2 liquid propellant rocket engine manages to produce
around 70,000 lbf of thrust which is enough to approach Mach 3 – 4. Anticipated body design of X-69 will
be similar to that of X-15 and SpaceshipTwo with additional concentration on its body material and its
efficiency for re-entry. Rocketmotor Two is an advanced powerplant that uses liquid/solid hybrid propellant
which can reduce weight unlike X-15’s XLR99. Again mission requirements can vary after entering to the
space based on altitude from earth, position of damaged satellite or position of undocking on-board
satellites.
An efficient re-entry system will allow X-69 simply glide and land back to base. This considers
advanced aerodynamics characteristics and wing, tail and body parts along with its material, preferably a
carbon composite.
6. 4.0 Conclusion and Recommendations:
4.1 Conclusion:
X-69 will make satellite deliveries a lot easier and cost effective as compared to existing methods.
Advanced technologies in electronics and computer have shrunk all the devices and have made them
compact. This gives rise to high market for small-sized satellites or Cubesats. Using a space-planes like X-
69 or an existing SpaceshipTwo makes the system efficient. Also we don’t have to pollute space by
disposing fairly working satellites. They can be brought back to earth, rework on it, modify the design and
send it back to space using X-69. X-69 will be an efficient glider that will use its aerodynamics to reach the
destination and safe landing.
4.2 Recommendations:
Air launch has many developments since X-15 launched from B-52 Stratofortress and other missile
launches from fighter jets. Motherships should have efficient performance to maintain launch altitude and
launch speed. Motherships are expected to be sophisticated from structural point of view. Air launches are
quite delicate while cruising with high speeds which might be vulnerable to structural integrity. Virgin
Galactic uses central air launch mechanism for Spaceship Two.
5.0 References:
https://www.globalaircraft.org/planes/x-15_hyper.pl
https://en.wikipedia.org/wiki/North_American_X-15
http://www.mach25media.com/Resources/X15FlightLog.pdf
http://er.jsc.nasa.gov/seh/ANASAGUIDETOENGINES%5B1%5D.pdf
https://www.nasa.gov/centers/armstrong/news/FactSheets/FS-052-DFRC.html
http://www.space.com/30245-x37b-military-space-plane-100-days.html
http://www.af.mil/AboutUs/FactSheets/Display/tabid/224/Article/104539/x-37b-orbital-test-vehicle.aspx
http://spaceflight101.com/spacecraft/x-37b-otv
https://www.nasa.gov/centers/armstrong/news/FactSheets/FS-052-DFRC.html
http://er.jsc.nasa.gov/seh/ANASAGUIDETOENGINES%5B1%5D.pdf
http://www.ijee.ie/articles/Vol13-4/ijee950.pdf
https://en.wikipedia.org/wiki/Boeing_X-37
7. 1
Configuration Design of X-69 CargoSat
AE 271 – Aircraft Design
Dr. Nikos Mourtos
Rushikesh Badgujar
San Jose State University
Charles W. Davidson College of Engineering
Aerospace Engineering
8. 2
Table of Contents
1. Introduction:..........................................................................................................................................3
2. Comparative Study: ..............................................................................................................................3
2.1 Configuration Comparison of Similar Airplanes:...............................................................................4
2.2 Discussion:..........................................................................................................................................7
3. Configuration Selection:...........................................................................................................................8
3.2 Wing Configuration:...........................................................................................................................9
3.3 Empennage Configuration: ...............................................................................................................11
3.4 Integration of the Propulsion system: ...............................................................................................11
3.5 Landing Gear Disposition:................................................................................................................11
3.6 Proposed configuration:....................................................................................................................12
4. References...............................................................................................................................................15
9. 3
1. Introduction:
This report describes the configuration design for X-69. Based on mission requirements, it is
necessary to propose a preliminary design which considers several aspects of aircraft like its general
characteristics, overall configuration, wing configuration, propulsion system, landing gear disposition,
etc. For initial guess, above parameters are referred to designs of similar aircrafts. X-planes like X-37B,
VG Spaceship One, VG Spaceship Two, Boeing X-20 Dyna Soar, X-15 are considered for comparative
study.
2. Comparative Study:
After more research, I want to add SpaceShip One for comparative study.
Table.1: Comparative study of similar airplanes
Parameters Boeing’s X-37B Virgin Galactic’s
Spaceship One
Virgin Galactic’s
Spaceship Two
Boeing’s X-20
Dyna-Soar
X-15
Crew none 1 pilot 2 crew and 6
passengers
1 pilot 1 pilot
Takeoff/Launch
weight
11,000 lb (4,990 kg) 21,428 lb (9,740 kg) 21,428 lb (9,740
kg)
11,387 lb (5,165
kg)
34,000 lb (15,420 kg)
Empty weight NA (electric
powered)
2,640 lb (1,200 kg) 15,000 lb (6,804
kg)
10,395 lb (4,715
kg)
14,600 lb (6,620 kg)
Thrust 157.4 lbf 700 kN) 16534.67 lbf (74 kN) 60,000 lbf (270
kN)
72,000 lbf (323
kN)
70,400 lbf (313 kN)
Critical Speed, V-
cr
(Orbital) 17,426 mph
(28,440 km/h)
2,170 mph (3,518
km/hr)
2,500 mph (4,000
km/hr)
17,500 mph
(28,165 km/hr)
4,520 mph (7,274
km/h)
Range, R 675 days (longest
flight)
40 mi (65 km) Planned apogee of
110 km
Earth orbit 22,000
nm (40,700 km)
280 mi (450 km)
Wing Area, S 161.4 ft2
(15 m2
) 273.34 ft2
(25.4
m2
) (estimated)
345 ft2
(32 m2
) 200 ft2
(18.6 m2
)
Wing span, b 14 ft 11 in (4.5 m) 16 ft. 5 in (8.05 m) 27 ft (8.3 m) 20 ft 10 in (6.34
m)
22 ft 4 in (6.8 m)
Aspect Ratio, AR 1.6 2.667 (estimated) 1.256 2.486
Type of Payload Satellites Pilot Crew Crew Crew
Powerplant Gallium Arsenide
Solar Cells with Li-
Ion batteries
1x N2O/HTPB
SpaceDev Hybrid
rocket,
1x Rocket
motorTwo
liquid/solid hybrid
rocket engine
1x Transtage
rocket engine
1x Reaction Motors
XLR99-RM-2 liquid
propellant rocket
engine
10. 4
2.1 Configuration Comparison of Similar Airplanes:
a) Boeing’s X-37B:
Fig.1: All views of Boeing X-37B
b) Virgin Galactic’s Spaceship One:
Fig.2: Front view Fig.3: Top view
12. 6
Fig.6: Front view with stabs trimmed up
d) Boeing’s X-20 Dyna Soar:
Fig.7: All views of X-20
13. 7
e) X-15:
Fig. 8: All views of X-15A-2 with external fuel tank
2.2 Discussion:
Following are the parameters that can have major impact on the design of X-69 as briefly
discussed in report 1:
As we can notice many facts are common in these aircrafts. Almost all the airplanes are designed
to land back dealing with hypersonic speeds and gliding. Also they use motherships to air launch except
X-37B which was launched using traditional rockets. Basically, a delta-wing pattern has been
implemented on almost all the above designs with certain variations. Delta-wings give efficient
performance at hypersonic speeds with better gliding as they descend.
14. 8
Low wing: X-37B, Spaceship Two, X-20 Dyna Soar
Advantages:
1. Over-wing exits.
2. While in space, it is quite easy to deploy small satellites from upper fuselage where wing does not
come in the way.
3. Easier to stick the main gear on.
4. Low wing doesn’t block any of the cabin.
5. Easy to access for maintenance and refueling.
Med Wing: X-15A-2. (also its predecessor, X-15A)
Advantages:
1. Med Wing provide best maneuverability.
2. Wing can be continuous through the fuselage.
3. Maintains structural integrity with the fuselage.
High Wing: Spaceship One
Advantages:
1. Quick loading and unloading.
2. Higher clearance from the ground providing less ground effect.
A certain disadvantage of high wing has been documented especially for Spaceship One. The
design was susceptible to roll excursions. It has been noticed that wind shear causes a large roll
immediately after ignition progressing into multiple rapid rolls. Although as it gains high speed upon
climb, this anomaly mitigates making the flight stable.
3. Configuration Selection:
3.1 Overall Configuration:
X-69 will be a land based aircraft.
A conventional type with Stabs and elevons.
Fuselage Configuration:
Spaceship Two and Spaceship One are built to accommodate 2 pilot and 6 – 8 crew with all facilities to
deal with gravitational variations while climbing, in space and while descending. Design of X-69 fuselage
will be conventional that seeks to accommodate satellites with sophisticated mechanisms to deploy or
undock satellites into LEO or dock back returning satellites without any damage to either X-69 or
satellite. Deployment systems like NanoRacks CubeSat Deployer (NRCSD) or XPOD Separation System
can be used based on the layout of racks and fuselage design. Direction of deployment can be from
sideways or rear since wing can be moved up or down with relatively less effect on X-69 maneuvering.
15. 9
3.2 Wing Configuration:
These type of planes have relatively different wing patterns unlike traditional airplanes. Wing
design of both Spaceship One and Two are standard and similar consisting Elevons and Stabilator.
Based on advantages of low wing position, X-69 will have Low wing.
Wing will be aft swept with certain angle.
This design will not necessarily require winglets due to presence of elevons and stabilators.
The horizontal stabilizer design for X-69 may lie between dual tail and twin tail.
Similarly, vertical tails will be connected and operated electrically on each tip of the horizontal tail to
control yaw and roll.
Elevons:
Like traditional airplanes, elevons are aircraft control surfaces that combine the functions of
elevator that controls pitch and aileron that controls roll. For above aircrafts, elevons are located
behind the stabilator (also called as stab) directly connected to the stick in the cockpit using cables.
Stabilator:
Stabilator, also called as stab, is fully movable aircraft stabilizer. Besides its usual function to
stabilize longitudinally, it is very useful device at high Mach number for changing the aircraft balance
within wide limits and for mastering the stick forces. In case of X-69, stab will control pitch when
trimmed up or down using electromechanical device and will control roll when moved independently.
Airfoils used for Wings:
X-69 like other X-planes does not need much of aerodynamics while climbing from around
45,000 ft to 50,000 ft. The climb is solely governed by rocket motor which takes barely 10-15
minutes to reach LEO. Ion Thrusters can be installed for efficient maneuver while in space.
Efficient airfoil selection will play vital role in returning phase and re-entry. X-69 will glide as it
descends after re-entry with required loitering to decelerate followed by landing approach. In the first
test flight of Spaceship One, landing procedure used modified version of a standard engine out
approach that is generally used by the military. HS 130 airfoil popular for dynamic soaring can be
used for wing to obtain efficient glide. From the research it is found that HS 130 delivers very less
drag and has characteristics of slope soaring. This airfoil is also used for elevons with 25% chord.
As shown in Fig.9 and Fig.10, the study has been done on HS 130 airfoil using Xfoil analyzing
pressure distribution, lift, drag and moment coefficients. Fig.10 considers various angle of attacks
from -2o
to 6o
at zero Mach for initial analysis.
16. 10
Fig.9: Cp at 6 deg of AoA with boundary layer
Fig.10: Pressure distribution Cp at various AoA’s
17. 11
3.3 Empennage Configuration:
As discussed earlier, the stabs and elevons will be controlled using electromechanical system.
Although, it is very important to pick the efficient configuration for better performance specially to glide,
loiter if necessary and safe landing. From the research so far, there are three options such as Tailplane
mounted, Twin tail boom or Wing mounted as shown in Fig. 11. For initial guess and referring to
previous designs, wing mounted configuration can be picked.
Fig.11: Empennage configurations
3.4 Integration of the Propulsion system:
Engine type: Rocket Motor engine
Engine Integration: Engine inside the fuselage from behind.
This is a X-type of plane that uses rocket motors for propulsion. For X-69, we seek to use solid/liquid
hybrid propellant rocket that can deliver thrust in the range of 60,000 lbf to 75,000 lbf. X-15 uses XLR99-
RM-2 liquid propellant rocket engine. Although this system makes it quite bulky to handle liquid
propellants and sloshing issues. On the other hand, Virgin Galactic and Scaled Composites used hybrid
rocket motor with benign fuel and oxidizer. The advantage of hybrid rocket motor is that it is controllable
and can be shut down at any time during boost phase of flight. It has less issues with sloshing.
As a new requirement, it seems necessary to use Reaction Control System(RCS) thrusters on
small scale for in-space maneuverability, attitude control to efficiently undock/dock satellites from X-69.
3.5 Landing Gear Disposition:
Landing Gear: Retractable gear
Nose-wheel landing gear
Landing gear integration: In the fuselage
Rear gears attached to wings retracting inwards.
18. 12
3.6 Proposed configuration:
Basic design of X-69 is almost similar to that of Spaceship Two. Although, based on
requirements such as, ground effects, type of propulsion system, docking/undocking mechanisms, design
of fuselage will be slightly different. As proposed earlier, fuselage will have sideway doors to deploy
satellites from upwards which is why low wing has been chosen. Placement of RCS thrusters has not
decided yet which is required for maneuver in space to conveniently operate docking/undocking.
Using Solidworks, following is the preliminary design of X-69 as shown in figures,
12,13,14,15,16 from various orientation and angles. Due to time constraints, the design is missing
complete tail which is of on-wing mounted type, proportionate size of fuselage, landing gears pockets for
retraction, etc. HS 130 airfoil has been used for wing and tail modelling.
Fig. 12: Isometric 3-D view of X-69
21. 15
4. References
(n.d.). Retrieved from
http://www.petervis.com/interests/published/Spaceshiptwo/Spaceshiptwo_Rocket.html.
(n.d.). Retrieved from http://www.nbcnews.com/storyline/virgin-voyage/how-spaceshiptwos-feathered-
wings-were-supposed-work-n240256.
(n.d.). Retrieved from http://bagera3005.deviantart.com/art/White-Knight-SpaceShip-One-158659309.
(n.d.). Retrieved from https://en.wikipedia.org/wiki/SpaceShipOne.
Airfoil Database. (n.d.). Retrieved from https://www.aerodesign.de/english/profile/profile_s.htm.
Airfoils and Airflow. (n.d.). Retrieved from https://www.av8n.com/how/htm/airfoils.html.
Donald Greer, hamory, P., Krake, K., & Drela, M. (n.d.). Design and Predictions for a High-Altitude
(Low-Reynolds-Number) Aerodynamic Flight Experiment.
Evans, M. (2013). The X-15 Rocket Planes. In M. Evans, The X-15 Rocket Planes.
FAA. (n.d.). Aerodynamics of Flight. In Gliding Flight Handbook.
Flight Training Center. (n.d.). Retrieved from http://flighttrainingcenters.com/training-aids/multi-
engine/engine-out-procedures/.
Global Aircraft. (n.d.). Retrieved from https://www.globalaircraft.org/planes/x-15_hyper.pl.
https://en.wikipedia.org/wiki/North_American_X-15. (n.d.). Retrieved from
https://en.wikipedia.org/wiki/North_American_X-15.
NASA. (n.d.). A NASA Guide to engines.
NASA factsheet. (n.d.). Retrieved from https://www.nasa.gov/centers/armstrong/news/FactSheets/FS-052-
DFRC.html.
Scaled Composites. (n.d.). Retrieved from http://www.scaled.com/projects/tierone/.
Space Flight Laboratory. (n.d.). Retrieved from http://utias-sfl.net/?page_id=87.
Space.com. (n.d.). Retrieved from http://www.space.com/30245-x37b-military-space-plane-100-
days.html.
Virgin Galactic. (n.d.). Retrieved from http://www.virgingalactic.com/human-spaceflight/our-vehicles/.
Virgin galactic fact sheet. (n.d.). Retrieved from
http://www.galacticexperiencesbydeprez.com/pdf/vg_vehicles_fact_sheet101411.pdf.
wired.com. (n.d.). Retrieved from https://www.wired.com/2010/10/test-pilot-describes-first-glide-flight-
of-spaceshiptwo/.
22. 1
Weight Sizing and Weight Sensitivities of X-69 CargoSat
AE 271 – Aircraft Design
Dr. Nikos Mourtos
Rushikesh Badgujar
San Jose State University
Charles W. Davidson College of Engineering
Aerospace Engineering
23. 2
Index
1. Introduction 3
2. Mission Weight Estimates 3
2.1.Database for Takeoff Weights and Empty Weights of Similar Airplanes 3
2.2.Determination of Regression Coefficients A and B 3
2.3. Determination of Mission Weights 5
2.3.1. Manual Calculation of Mission Weights 6
2.3.2. Calculation of Mission weights using the AAA program 9
3. Takeoff Weight Sensitivities 12
3.1.Manual Calculation of Takeoff Weight Sensitivities 12
3.2.Calculation of Takeoff Weight Sensitivities using the AAA program 16
3.3. Trade studies 17
4. Discussion 19
5. Conclusions and Recommendations 21
5.1.Conclusions 21
5.2.Recommendations 21
6. References 22
7. Appendices 23
24. 3
1. Introduction:
The mission profile of X-69 divides its flight into two phases. X-69 makes an air launch and climbs at
supersonic speed in Phase I. This phase will last for about 90 seconds in which X-69 will reach to altitude
of about 360,000 ft (110 km) at Mach 3.0 – 3.5. Deploying the payload (satellites) into space and docking
back returning satellites if required by mission, it will descend and land using its gliding characteristics
decelerating itself to subsonic speeds in Phase II. X-69 will have reaction control system such as RV-105
RCS Thruster block or Vernor Engine to control transcend between phase I and II. These thrusters use very
less fuel to maintain the required thrust in x, y and z direction for maneuvering and attitude control. In phase
I for weight analysis, we consider supersonic cruise fuel-fraction for initial calculations. In phase II for
weight analysis, we consider sail plane fuel-fraction which eventually is similar to small home built
airplanes.
2. Mission Weight Estimates:
2.1. Database for Takeoff Weights and Empty Weights of Similar Airplanes:
Table 2.1 provides a database for takeoff and empty weights of similar airplanes. Boeing X-37B is
an exception since this ROT vehicle is electric powered which has same empty weight as takeoff.
Table 2.1: Database for Takeoff weights and empty weights of Similar Airplanes
Airplane Type Takeoff Weight Empty Weight
Lockheed CL-1200 Lancer Supersonic 35,000 lbs (15,900 kg) 17,885 lbs (8,112 kg)
Martin Marietta X-24B Supersonic 13,800 lbs (6,260 kg) 8,500 lbs (3,855 kg)
Virgin Galactic Spaceship One Supersonic 10,560 lbs (4,800 kg) 2,640 lbs (1,200 kg)
Virgin Galactic Spaceship Two Supersonic 21,428 lbs (9,740 kg) 10,423 lbs (4,272.8 kg)
Boeing X-20 Dyna Soar Supersonic 11,387 lbs (5,165 kg) 10,395 lbs (4,715 kg)
North American X-15 Supersonic 34,000 lbs (15,420 kg) 14,600 lbs (6,620 kg)
Lockheed Martin X-33 suborbital spaceplane 285,000 lbs (129,000 kg) 75,000 lbs (34,019.43 kg)
NASA X-38 CRV re-entry vehicle 54,500 lbs (24,721 kg) 23,500 lbs (10,660 kg)
Douglas X-3 Stiletto Supersonic 23,840 lbs (10,810 kg) 16,120 lbs (7,310 kg)
Ryan X-13 Vertijet VTOL jet aircraft 7,200 lbs (3,272 kg) 5,334 lbs (2,424 kg)
Boeing X-37B Reusable Orbital Test vehicle 11,000 lbs (4,990 kg) 11,000 lbs (4,990 kg)
North American X-10 cruise missile 42,300 lbs (19,187 kg) 25,800 lbs (11,703 kg)
Boeing X-40 Reusable launch vehicle 3,700 lbs (1,640 kg) 2,500 lbs (1,100 kg)
2.2. Determination of Regression Coefficients A and B:
Before initiating the calculation for determination of mission weights and empty weight, it
is necessary to make an initial guess for takeoff weight based on similar airplanes take off weight
data. Fig. 1 is the plot of empty weight v/s takeoff weight of similar airplanes as tabulated in table.1.
Initial takeoff weight can be guessed close to trend line. For X-69 the payload weight will be 14,000
lbs, similar to that of Spaceship Two.
25. 4
Fig. 1: Weight trends for Space-planes
Fig. 2: log-log plot of weight data
It is necessary to determine regression coefficients A and B for X-69. Fig.2 is a log-log plot of weight
data of similar airplanes that are considered. Equation of trend line will give regression coefficients A and
B as follows:
Equation of trend line is:
𝑦 = 0.8613. 𝑥 + 0.3651
In this case,
y = log10(WE)
x = log10(WTO)
X-69 CargoSat
0
5000
10000
15000
20000
25000
30000
0 10000 20000 30000 40000 50000 60000
EmptyWeight,WE,lbs
Gross Take-off weight,WTO , lbs
Weight Trends for Space-planes
X-69
y = 0.8613x + 0.3651
3
3.2
3.4
3.6
3.8
4
4.2
4.4
4.6
3 4 5
log10(WE)
log10(WTO)
log10(WE) v/s log10(WTO)
26. 5
from equation 2.16 in Roskam,
𝑊𝐸 = 10
{
log(10) 𝑊 𝑇𝑂−𝐴
𝐵
}
Simplifying above equation, we get
Log10(𝑊𝐸) =
1
𝐵
log10 𝑊𝑇𝑂 −
𝐴
𝐵
Therefore,
1
𝐵
= 0.8613
𝐵 = 1.161
And
−
𝐴
𝐵
= 0.3651
𝐴 = −0.424
Hence we find the regression coefficients as A = -0.424 and B = 1.161.
2.3. Determination of Mission Weights:
There are two methods such as manual calculation and using AAA program to determine
mission weights. Both methods begin by guessing a takeoff weight followed by sequential phases
of flight like engine start and warmup, taxi, takeoff, climb, loiter if necessary, descent and
approach and landing. The methods are described as follows:
𝑊𝑡𝑎𝑘𝑒𝑜𝑓𝑓 = 32,000 𝑙𝑏𝑠 (𝐺𝑢𝑒𝑠𝑠)
As discussed, X-69 flight consists of two phases in its complete flight. It will make an air
launch from mothership with a clean release and climb at supersonic speed. Although while
descending and landing X-69 will glide decelerating to subsonic speed using aerodynamics and
efficient wing configuration. Hence fuel-fractions will be considered according to the mission
phase of the flight. Subsonic glide fuel fractions are average of fuel-fractions of light weight
aircrafts like Homebuilt, Single Engine, Twin Engine and Agricultural airplanes. Highlighted
section of table 2.2 are the fuel-fractions considered for two phases of X-69 flight and hence to
calculate its takeoff weight.
Referring to table 2.1. Suggested Fuel-fractions for Several Mission Phases in Roskam
book,
Table 2.2. Suggested Fuel-fractions for Several Mission Phases
Mission Phase Engine start Takeoff/Air launch Climb Descent Landing, Taxi and Shutdown
Supersonic Cruise 0.990 0.995 0.92-0.87 0.985 0.992
Subsonic Glide 0.995 0.997 0.994 0.993 0.995
Referring to Table 2.15. Regression Line Constants A and B in Roskam, regression
coefficients A and B are 0.4221 and 0.9876 respectively since X-69 will be in supersonic flight.
27. 6
2.3.1.Manual Calculation of Mission Weights:
The fuel-fraction, mff for each phase is defined as the ratio of end weight to begin weight.
The next step is to assign a numerical value to the fuel-fraction corresponding to each mission
phase. This is done as follows referring to table. 2.2:
Phase 1. Engine Start and warm up:
Begin weight is WTO. End weight is W1. The fuel fraction for this phase is by definition
given by: W1/WTO.
Therefore,
𝑓𝑢𝑒𝑙 − 𝑓𝑟𝑎𝑐𝑡𝑖𝑜𝑛 =
𝑊1
𝑊𝑇𝑂
0.990 =
𝑊1
32,000 𝑙𝑏𝑠
𝑊1 = 31,680 𝑙𝑏𝑠
Phase 2. Clean Release, Air launch/Take off:
Begin weight is W1. End weight is W2. The fuel fraction for this phase is W2/W1.
Therefore,
𝑓𝑢𝑒𝑙 − 𝑓𝑟𝑎𝑐𝑡𝑖𝑜𝑛 =
𝑊2
𝑊1
0.995 =
𝑊2
31,680 𝑙𝑏𝑠
𝑊2 = 31,521.6 𝑙𝑏𝑠
Phase 3. Climb at Supersonic speed:
Begin weight is W2. End weight is W3. The fuel fraction for this phase is W3/W2.
To calculate fuel-fraction for climb, we need to know following parameters:
Change in altitude, ∆ℎ
Rate of climb
L/D ratio
Specific fuel consumption, cj
Therefore,
𝑓𝑢𝑒𝑙 − 𝑓𝑟𝑎𝑐𝑡𝑖𝑜𝑛 =
𝑊3
𝑊2
0.9951 =
𝑊3
31,521.6 𝑙𝑏𝑠
28. 7
𝑊3 = 31367.14 𝑙𝑏𝑠
Phase 4. Drop from the space and descent towards earth from 80,000 ft:
Begin weight is W4. End weight is W3. The fuel fraction for this phase is W4/W3.
Therefore,
𝑓𝑢𝑒𝑙 − 𝑓𝑟𝑎𝑐𝑡𝑖𝑜𝑛 =
𝑊4
𝑊3
0.993 =
𝑊4
31,367.14 𝑙𝑏𝑠
𝑊4 = 31,147.57 𝑙𝑏𝑠
Phase 5. Approach, Landing, Taxi and shutdown:
Begin weight is W5. End weight is W4. The fuel fraction for this phase is W5/W4.
Therefore,
𝑓𝑢𝑒𝑙 − 𝑓𝑟𝑎𝑐𝑡𝑖𝑜𝑛 =
𝑊5
𝑊4
0.995 =
𝑊5
31,147.57 𝑙𝑏𝑠
𝑊5 = 30,991.84 𝑙𝑏𝑠
Therefore, the mission fuel-fraction, Mff is given as:
𝑀𝑓𝑓 = (
𝑊1
𝑊𝑇𝑂
) ∗ ∏ (
𝑊𝑖+1
𝑊𝑖
)
𝑖=5
𝑖=1
𝑀𝑓𝑓 = (
𝑊1
𝑊𝑇𝑂
) (
𝑊2
𝑊1
) (
𝑊3
𝑊2
) (
𝑊4
𝑊3
) (
𝑊5
𝑊4
)
𝑀𝑓𝑓 = (
𝑊5
𝑊𝑇𝑂
)
𝑀𝑓𝑓 = (
30,991.84
32,000
)
𝑀𝑓𝑓 = 0.9685
Weight of fuel used, Wf_used
𝑊𝐹 𝑢𝑠𝑒𝑑
= (1 − 𝑀𝑓𝑓)𝑊𝑇𝑂
𝑊𝐹 𝑢𝑠𝑒𝑑
= (1 − 0.9685) × 32,000
29. 8
𝑊𝐹 𝑢𝑠𝑒𝑑
= 1,008 𝑙𝑏𝑠
A = -0.424, B = 1.161
𝑊𝐸 = 10(log10 𝑊 𝑇𝑂−𝐴)/𝐵
𝑊𝐸 = 10
log10 32,000+0.424
1.161
𝑊𝐸 = 17,603.8 𝑙𝑏𝑠
A tentative value for WOE is found from equation below:
𝑊𝑂𝐸_𝑡𝑒𝑛𝑡 = 𝑊𝑇𝑂 − 𝑊𝐹 𝑢𝑠𝑒𝑑
− 𝑊𝑃𝐿
Payload weight, WPL is 14,000 lbs.
𝑊𝑂𝐸_𝑡𝑒𝑛𝑡 = 32,000 − 1,008 − 14,000
𝑊𝑂𝐸_𝑡𝑒𝑛𝑡 = 16,992 𝑙𝑏𝑠
A tentative value for WE is found from equation below:
𝑊𝐸𝑡𝑒𝑛𝑡
= 𝑊𝑂𝐸𝑡𝑒𝑛𝑡
− 𝑊𝑇𝐹𝑂 − 𝑊𝑐𝑟𝑒𝑤
𝑊𝑇𝐹𝑂 = 0.005 × 𝑊𝑇𝑂 = 0.005 × 32,000
𝑊𝑇𝐹𝑂 = 160 𝑙𝑏𝑠
𝑊𝐸𝑡𝑒𝑛𝑡
= 16,992 − 160 − 350
𝑊𝐸𝑡𝑒𝑛𝑡
= 16482 𝑙𝑏𝑠
Comparing WE-tent and WE-allowable/ WE,
|𝑊 𝐸−𝑊 𝐸 𝑡𝑒𝑛𝑡|
𝑊 𝐸+𝑊 𝐸 𝑡𝑒𝑛𝑡
2
=
|17,603.8−16482|
17,603.8+16482
2
× 100 = 6.583%
Using MATLAB, WTO can be iterated to obtain required comparison less than 0.5%. The code can
be referred from Appendix C.
After iterating, WTO = 34,200 lbs gives comparison percentage of about 0.214% which is less than
0.5% with empty weight, WE = 18,641.42 lbs.
Table. 3: mission weights with respect to selected takeoff weight = 34,200 lbs
Engine start and warmup, w1 33858.0 lbs
Air launch/ takeoff, w2 33688.7 lbs
Climb, w3 33523.6 lbs
Descent, w4 33289.0 lbs
Land and taxi, w5 33122.5 lbs
Weight of fuel used, Wf 1077.33 lbs
30. 9
2.3.2.Calculation of Mission weights using the AAA program:
Before starting the calculation for take-off weight, it is necessary to set up the configuration of X-
69 in the software. Appendix D shows initial steps to set the parameters and configuration of aircraft:
In AAA program, after configuring X-69, we start with weight analysis by defining mission profile
and respective fuel-fractions. Fig. 3 shows sequentially arranged segment-wise mission profile with
required fuel-fractions:
Fig. 3. Mission profile of X-69 with fuel-fraction.
After mission profile is defined, we obtain regression coefficients based of empty and takeoff
weights of similar airplanes that are accounted in table.1. Clicking on “Weight Sizing” opens following
window as shown in Fig. 4 where user needs to feed in similar airplanes data as I did it for X-69.
It is necessary to maintain the airplane data less scattered. The more the airplanes, more will be the
accuracy for regression coefficients. For X-69, I found up to 11 similar airplanes that were considered to
compute regression coefficients and will be used to compute takeoff weight in next step. Fig. 5 is shows
the trend line for empty weight v/s takeoff weight from which regression coefficients were obtained.
32. 11
After obtaining regression coefficients A and B, it is safe to proceed for takeoff weight. We input
required parameters in “Take-off weight: Flight condition 1” window as shown in Fig. 6. We input same
regression coefficients A and B that we obtained in previous step. Any near-takeoff weight can be guessed
under WTOest. There are no passengers in X-69 but 2 pilots weighing approximately 175 lbs each. Rest of
the payload of satellites have been considered under Wcargo = 14,000 lbs. Fuel fraction of trapped fuel and
oil is assumed to be 0.005% as referred from Roskam. X-69 won’t need any reserve fuel, so Mres=0.
After hitting calculate, it gives following output parameters with slightly different value of takeoff
weight than that obtained from manual calculation.
Fig. 7 shows the design point for takeoff weight using the same equations that were used to perform
manual calculations. AAA program gives WTO as design point equal to 36505.2 lbs which is little higher
number than that obtained from manual calculation with WTO = 34,200 lbs.
Fig. 6. Takeoff weight: Flight condition 1.
33. 12
Fig. 7. Design point on trend line @ 36505.2 lbs
3. Takeoff Weight Sensitivities:
3.1. Manual Calculation of Takeoff Weight Sensitivities:
Before starting sensitivity calculations, it should be checked if equation 2.24 from Roskam
yields approximately same takeoff weight, WTO as we obtained from iterative method in section
2.3. To do that we can substitute values of regression coefficients A, B, C and D in equation 2.24
as stated below:
log10 𝑊𝑇𝑂 = 𝐴 + 𝐵 log10(𝐶. 𝑊𝑇𝑂 − 𝐷)
Substituting A=-0.424, B=1.161, C=0.9635 and D=14,350 lbs and using small matlab
solver from Appendix C, we get
log10 𝑊𝑇𝑂 = −0.424 + 1.161 × log10(0.9635 × 𝑊𝑇𝑂 − 14,350)
𝑊𝑇𝑂 = 34,280.4 𝑙𝑏𝑠
WTO that we just obtained is quite close to th`at we got from iterative method, hence we
can move ahead with this takeoff weight for sensitivity calculations.
X-69 will not be cruising at any time throughout its flight since it will climb at supersonic
speed as it drops from mothership. Hence there is no cruise consideration while calculating
sensitivities either in AAA program.
After preliminary sizing, it is mandatory to conduct sensitivity studies on parameters such as
34. 13
Payload, WPL:
Sensitivity of Take-off weight to Payload Weight:
From section 2.7.2 Sensitivity of Take-off weight to Payload weight of Roskam, sensitivity
of take-off weight to payload weight is given by:
𝜕𝑊𝑇𝑂 𝜕𝑊𝑃𝐿 = 𝐵. 𝑊𝑇𝑂/(𝐷 − 𝐶(1 − 𝐵). 𝑊𝑇𝑂)⁄
𝐴 = −0.424, 𝐵 = 1.161
𝐶 = {1 − (1 + 𝑀𝑟𝑒𝑠)(1 − 𝑀𝑓𝑓) − 𝑀𝑡𝑓𝑜}
𝑊𝐹𝑟𝑒𝑠
= 𝑀𝑟𝑒𝑠. (1 − 𝑀𝑓𝑓). 𝑊𝑇𝑂
𝑀𝑟𝑒𝑠 = (𝑊𝐹𝑟𝑒𝑠
/((1 − 𝑀𝑓𝑓). 𝑊𝑇𝑂))
No reserves, therefore
𝑀𝑟𝑒𝑠 = 0
𝐶 = {1 − (1 + 𝑀𝑟𝑒𝑠)(1 − 𝑀𝑓𝑓) − 𝑀𝑡𝑓𝑜}
𝐶 = {1 − (1 + 0)(1 − 0.9685) − 0.005}
𝐶 = 0.9635
𝐷 = 𝑊𝑃𝐿 + 𝑊𝑐𝑟𝑒𝑤
𝐷 = 14,000 + 350
𝐷 = 14,350 𝑙𝑏𝑠
𝜕𝑊𝑇𝑂 𝜕𝑊𝑃𝐿 = 𝐵. 𝑊𝑇𝑂/(𝐷 − 𝐶(1 − 𝐵). 𝑊𝑇𝑂)⁄
𝜕𝑊𝑇𝑂 𝜕𝑊𝑃𝐿 = 1.161 ×
34,280.4
14,350 − 0.9635(1 − 1.161) × 34,280.4
⁄
𝜕𝑊𝑇𝑂 𝜕𝑊𝑃𝐿 = 2.021⁄
This means that for each pound of payload added, the airplane take-off gross weight will
have to be increased by 2.021 lbs and is called growth factor due to payload for X-69.
Empty weight, WE
Sensitivity of Take-off weight to Payload Weight:
From section 2.7.3, Sensitivity of Take-off weight to Empty weight of Roskam, sensitivity
of take-off weight to payload weight is given by:
𝜕𝑊𝑇𝑂 𝜕𝑊𝐸 =
𝐵𝑊𝑇𝑂
[10{(log10 𝑊 𝑇𝑂−𝐴)/𝐵}]
⁄
𝜕𝑊𝑇𝑂 𝜕𝑊𝐸 = 1.161 ×
34,280.4
[10{(log10 34,280.4+0.424)/1.161}]
⁄
𝜕𝑊𝑇𝑂 𝜕𝑊𝐸 = 2.13⁄
For each lb of increase in empty weight, the take-off weight will increase by 2.13 lbs and
is a growth factor due to empty weight for X-69.
35. 14
Range, R
Sensitivity of Take-off weight to Range:
Estimated range of X-69 is 120nm (110 km) return trip including re-entry and landing. X-
69 will climb at supersonic speed and hence the characteristics for calculating sensitivities of take-
off weight and range will be similar to that of fighter airplanes. From It is necessary to calculate a
factor F using equation 2.44 in Roskam as given below:
𝐹 =
−𝐵. 𝑊𝑇𝑂
2
{𝐶𝑊𝑇𝑂. (1 − 𝐵) − 𝐷}
× (1 + 𝑀𝑟𝑒𝑠)𝑀𝑓𝑓
Substituting values of B, C, D in above equation.
𝐹 =
−1.161 × 34,280.42
{0.9635 × 34,280.4 × (1 − 1.161) − 14,350}
× 0.9685
𝐹 = 67184.66
𝜕𝑊𝑇𝑂
𝜕𝑅
=
𝐹𝑐𝑗
𝑉𝐿
𝐷
considering cruise out numbers for X-69 from Roskam,
𝑐𝑗 = 1.25, 𝑉 = 459 𝑘𝑡𝑠 (𝑠𝑢𝑝𝑒𝑟𝑠𝑜𝑛𝑖𝑐, 3186
𝑘𝑚
ℎ𝑟
= 𝑀𝑎𝑐ℎ 3.0) ,
𝐿
𝐷
= 7: 1
Therefore,
𝜕𝑊𝑇𝑂
𝜕𝑅
=
67184.66 × 1.25
3186 𝑘𝑚/ℎ𝑟 × 7
𝜕𝑊𝑇𝑂
𝜕𝑅
= 3.766 𝑙𝑏𝑠/𝑘𝑚
Hence for every increase of in kilometer, gross take-off weight will increase by 3.766 lbs.
Endurance, E
Sensitivity of Take-off weight to Endurance:
Same as Range, sensitivity of takeoff weight to endurance is given by
𝜕𝑊𝑇𝑂
𝜕𝐸
=
𝐹𝑐𝑗
𝐿
𝐷
𝜕𝑊𝑇𝑂
𝜕𝐸
=
67184.66 × 1.25
7
𝜕𝑊𝑇𝑂
𝜕𝐸
= 11997.26 𝑙𝑏𝑠/ℎ𝑟
36. 15
Lift-to-drag ratio, L/D
Sensitivity of Take-off weight to Lift-to-drag ratio with respect to range requirement:
𝜕𝑊𝑇𝑂
𝜕 (
𝐿
𝐷)
= −
𝐹𝑅𝑐𝑗
𝑉 (
𝐿
𝐷)
2
𝜕𝑊𝑇𝑂
𝜕 (
𝐿
𝐷)
= −
67184.66 × 94.488 𝑘𝑚 × 1.25
3186 𝑘𝑚/ℎ𝑟 × 72
X-69 will use fuel only while climbing. As said earlier, it will just glide while descending
without any use of fuel. Hence range, R in above equation is taken only when it is climbing from
50,000 ft to 360,000 ft which gives 94.488 km of climb.
Hence,
𝜕𝑊𝑇𝑂
𝜕 (
𝐿
𝐷)
= −50.83 𝑙𝑏𝑠
If the lift-to-drag ratio of the airplane were 16 instead of the assumed 14, the design take-
off gross weight would decrease by 16-14=2× 50.83=101.66 lbs.
Specific fuel consumption, cj
Sensitivity of Take-off weight to specific fuel consumption, cj with respect to range
requirement:
𝜕𝑊𝑇𝑂
𝜕𝑐𝑗
=
𝐹𝑅
𝑉
𝐿
𝐷
𝜕𝑊𝑇𝑂
𝜕𝑐𝑗
=
67184.66 × 94.488 𝑘𝑚
3186 𝑘𝑚/ℎ𝑟 × 7
𝜕𝑊𝑇𝑂
𝜕𝑐𝑗
= 284.64
𝑙𝑏𝑠
𝑙𝑏𝑠
/(
𝑙𝑏𝑠
ℎ𝑟
)
If specific fuel consumption was incorrectly assumed to be 0.5 and in reality turns out to
be 0.9, the design take-off gross weight will increase by 0.9-0.5=0.4× 284.64=113.86 lbs.
3.2. Calculation of Takeoff Weight Sensitivities using the AAA program:
AAA program provides sensitivity computation if required parameters have been inserted. If
sequential procedure is followed, i.e starting with mission profile, obtaining regression coefficients and
hence takeoff weight, then sensitivity automatically considers those basic parameters. Hitting calculate
button gives sensitivities of takeoff weight with payload, empty weight, range, specific fuel
consumption, L/D ratio and endurance.
37. 16
Since X-69 will never cruise neither loiter, there is no takeoff sensitivity with respect to range. As
stated earlier, after completing its mission in space it will be dropped towards earth with accurate
attitude using reaction control system. As it reaches 80,000 – 85,000 ft, it will use its wing and
aerodynamics to slow down and glide to destination. Although its total range will be 120 nm
considering return trip excluding cruising, loitering phase of flight.
Fig. 8. Sensitivity calculation and growth factors.
There is significant percentage difference between manually calculated and AAA computed
sensitivities. Regression coefficients A and B are the major cause responsible for this difference. It is
unclear that what method does AAA program use to calculate regression coefficients A and B using same
similar airplane database and hence trend line that is been used for manual calculation. Although from the
linear characteristics and using method described in Roskam Book, we obtain different A and B for manual
calculations than from AAA program. Table. 5 describes the percentage difference which on an average is
up to 30%.
3.3. Trade Studies:
As discussed earlier, X-69 will never cruise or loiter in its flight. Hence trade studies of Range, R
and other parameters were not performed. Although, trade studies were performed for other significant
parameters that are required for X-69 such as specific fuel consumption, rate of climb, payload, etc.
As asked, while performing trade studies, takeoff weight has been kept constant throughout. AAA
program is used for trade studies. Table. 4 shows the trade study. Certain amount of payload has to be
traded-off for increase specific fuel consumption to accommodate more fuel at certain climb rate keeping
the takeoff weight constant. Also we can see that if X-69 needs to climb faster, specific fuel consumption
has to be increased reducing the payload weight.
Fig. 9 shows the plot of trade study. If X-69 seeks for higher specific fuel consumption, payload
has to be reduced depending on the climb rate.
This method can be used to read above plot and is explained below:
38. 17
For instance, say X-69 is climbing at rate of 150,000 ft/min with payload of 14,000 lbs, from plot,
it will need 0.9 lb/hr/lb of fuel shown by blue line in Fig. 10. Now if owner wants X-69 to climb at higher
rate, say 200,000 ft/min keeping the same payload of 14,000 lbs and takeoff weight which is already
constant, it will have to consume approximately 1.275 lb/hr/lb of fuel shown by red line in Fig. 10. At the
same time, if owner changes his mind and wants to reduce climb rate to 150,000 ft/min back again then
payload has to be reduced to approximately 13,925 lbs gaining more specific fuel consumption.
Table. 6: Trade study of payload with specific fuel consumption and rate of climb
rate of climb, ft/min 150000 200000 250000 300000
Takeoff weight, WTO, lbs specific fuel consumption, cj, lb/hr/lb Payload weight, WPL, lbs
36505.1 0.6 14050.35 14077.15 14089.6 14099.3
36505.3 0.7 14039.9 14065.35 14079.75 14091.15
36505.1 0.8 14017.5 14053.4 14069.85 14083.25
36505.1 0.9 14001.15 14041.5 14060 14074.8
36505.1 1 13984.9 14029.65 14050.25 14066.6
36505.1 1.1 13968.57 14017.75 14040.5 14058.35
36505.2 1.2 13952.15 14005.95 14030.55 14050.3
36505 1.3 13935.9 13994 14020.75 14042
36505.2 1.4 13920.55 13982.2 14011 14033.9
36505.2 1.5 13903.3 13970.35 14001.15 14025.65
Fig. 9: Trade study of payload with specific fuel consumption and rate of climb
Design Point
13850
13900
13950
14000
14050
14100
14150
0.5 0.7 0.9 1.1 1.3 1.5 1.7
payloadweight,WPL,lbs
specific fuel consumption, cj , lb/hr/lb
Trade study of payload weight vs specific fuel
consumption with takeoff weight constant
150000 ft/min 200000 ft/min 250000 ft/min 300000 ft/min
39. 18
Fig.10: Method example to read the trade study plot.
Red dot in fig. 9 can be good design point where we get to increase payload by around 25 lbs
maintaining climb rate of 200,000 ft/min but trading off fuel consumption by around 0.15 lb/hr/lb which is
not much since climb time is hardly 2 minutes. At this design point, X-69 will travel at Mach 3.0.
4. Discussion:
AAA software refers the theory of Roskam book up to some extent. Although, it is unclear that how
diverse airplanes data it can handle that should give close values to that obtained from manual
calculation. As we can observe that there is slight difference in final parameters in AAA program than
we obtained using manual calculations. Following is table. 5 showing percentage difference between
manually calculated and AAA program computed parameters:
Table. 5: Percentage difference for Mission weights in lbs
Parameters Manually Calculated AAA program computed % difference
Takeoff weight, WTO 34,200 36505.2 6.315%
Empty Weight, WE 18,641.42 21028.8 11.35%
Fuel weight, WF 1077.33 1124.6 4.2%
Assuming that AAA program estimation has higher accuracy than manual calculation, it is
beneficial to proceed with AAA computed takeoff weight for trade studies. Also the fuel-fractions
considered for each phase in both manual calculation and AAA program are same. Takeoff weight and
hence other dependent parameters also depend on regression coefficients A and B. And as we can see those
coefficients are slightly different giving rise difference in takeoff weight and other parameters.
Same as mission weight calculation, we have considerable difference in parameters obtained from
manual calculation and using AAA program. Table. 6 gives the percentage difference between significant
parameters:
40. 19
Table. 6: Percentage difference for sensitivities
Parameters Manually calculated AAA computed % difference
𝝏𝑾 𝑻𝑶 𝝏𝑾 𝑷𝑳⁄ 2.021 2.84 28.84%
𝝏𝑾 𝑻𝑶 𝝏𝑾 𝑬⁄ 2.13 1.62 31.5%
𝝏𝑾 𝑻𝑶 𝝏𝑬⁄ 11997.26 17925.5 33.07%
𝝏𝑾 𝑻𝑶 𝝏( 𝑳
𝑫⁄ )⁄ -50.83 -60.1 2.995%
𝝏𝑾 𝑻𝑶 𝝏𝒄𝒋⁄ 284.64 336.5 15.41%
Takeoff weight is very significant and primary parameter for aircraft design. Hence it is necessary
and mandatory to analyze takeoff weight sensitivities with respect to payload weight, empty weight,
endurance, range, life-to-drag L/D ratio and specific fuel consumption (sfc).
Considering the accuracy of AAA program, it is preferable to consider the computed values for
design. Hence for each pound of payload added, the airplane take-off gross weight will have to be increased
by 2.84 lbs and is called growth factor due to payload for X-69. Similarly, for each lb of increase in empty
weight, the take-off weight will increase by 1.62 lbs and is a growth factor due to empty weight for X-69.
If the lift-to-drag ratio of the airplane were 16 instead of the assumed 14, the design take-off gross weight
would decrease by 16-14=2× 60.1=120.2 lbs. If specific fuel consumption was incorrectly assumed to be
0.5 and in reality turns out to be 0.9, the design take-off gross weight will increase by 0.9-0.5=0.4×
336.5=134.6 lbs.
Observing takeoff weight sensitivity for specific fuel consumption, cj and trade study plot fig. 9,
we can see that if X-69 owner desires to have more payload, specific fuel consumption has to be sacrificed
making the engine less efficient per pounds of force. Although keeping the takeoff weight constant and
varying payload has less impact on specific fuel consumption as compared that with takeoff weight.
Referring the table. 6, we can see that for slight reduction in specific fuel consumption, cj gross takeoff
weight increases drastically.
41. 20
5. Conclusion and Recommendations:
5.1. Conclusions:
X-69 is desired to have the payload weight of 14,000 lbs that will accommodate satellites, racks to
place and hold satellites and instruments together unharmed from high speed climb of X-69. Having
mentioned the consistency of payload weight, takeoff weight of X-69 has been calculated using manual
calculation as well as AAA program. After considering all parametric aspects, 36,505 lbs will be a design
point for takeoff weight of X-69. X-69 will be manufactured using composites hence regression coefficients
A and B have to be calculated separately for manual calculation as mentioned in Roskam. After doing trade
study, the optimal design point is estimated to be 200,000 ft/min of climb rate at Mach 3.0 with specific
fuel consumption of 1.1 lb/hr/lb at payload of about 14025 lbs trading off 0.15 lb/hr/lb of fuel consumption.
X-69 will be using hybrid rocket motor with nylon as solid fuel and liquid nitrous oxide as liquid oxidizer.
5.2. Recommendations:
The weight analysis has been done very diverse data obtained from various resources. The
compared similar airplanes have very diverse configurations with respect to their missions. Hence unlike
conventional airplanes, there is no reference on previously done analysis on these type of airplanes in
Roskam or any other resources. Altitude limit can be extended to low earth orbit to about 160 km. Also if
reaction control thrusters are efficient enough, besides maneuvering for satellites they can be used to orbit
X-69 over certain location before it drops to earth gravity.
From initial research, HS 130 airfoil will be used for wing, elevons and stabs since it has high
gliding efficiency at high altitude and speeds.
42. 21
6. References
(n.d.). Retrieved from
http://www.petervis.com/interests/published/Spaceshiptwo/Spaceshiptwo_Rocket.html.
(n.d.). Retrieved from http://www.nbcnews.com/storyline/virgin-voyage/how-spaceshiptwos-feathered-
wings-were-supposed-work-n240256.
(n.d.). Retrieved from http://bagera3005.deviantart.com/art/White-Knight-SpaceShip-One-158659309.
(n.d.). Retrieved from https://en.wikipedia.org/wiki/SpaceShipOne.
Airfoil Database. (n.d.). Retrieved from https://www.aerodesign.de/english/profile/profile_s.htm.
Airfoils and Airflow. (n.d.). Retrieved from https://www.av8n.com/how/htm/airfoils.html.
Donald Greer, hamory, P., Krake, K., & Drela, M. (n.d.). Design and Predictions for a High-Altitude
(Low-Reynolds-Number) Aerodynamic Flight Experiment.
Evans, M. (2013). The X-15 Rocket Planes. In M. Evans, The X-15 Rocket Planes.
FAA. (n.d.). Aerodynamics of Flight. In Gliding Flight Handbook.
Flight Training Center. (n.d.). Retrieved from http://flighttrainingcenters.com/training-aids/multi-
engine/engine-out-procedures/.
Global Aircraft. (n.d.). Retrieved from https://www.globalaircraft.org/planes/x-15_hyper.pl.
https://en.wikipedia.org/wiki/North_American_X-15. (n.d.). Retrieved from
https://en.wikipedia.org/wiki/North_American_X-15.
NASA. (n.d.). A NASA Guide to engines.
NASA factsheet. (n.d.). Retrieved from https://www.nasa.gov/centers/armstrong/news/FactSheets/FS-052-
DFRC.html.
Scaled Composites. (n.d.). Retrieved from http://www.scaled.com/projects/tierone/.
Space Flight Laboratory. (n.d.). Retrieved from http://utias-sfl.net/?page_id=87.
Space.com. (n.d.). Retrieved from http://www.space.com/30245-x37b-military-space-plane-100-
days.html.
Virgin Galactic. (n.d.). Retrieved from http://www.virgingalactic.com/human-spaceflight/our-vehicles/.
Virgin galactic fact sheet. (n.d.). Retrieved from
http://www.galacticexperiencesbydeprez.com/pdf/vg_vehicles_fact_sheet101411.pdf.
wired.com. (n.d.). Retrieved from https://www.wired.com/2010/10/test-pilot-describes-first-glide-flight-
of-spaceshiptwo/.
43. 22
7. Appendices
7.1. Appendix A: Mission Requirements:
Table. 7: Mission requirement for X-69
Crew 2
Payload (Satellites), WPL 14,000 lbs
Take-off weight (WTO) calculated 30,800 lbs
Air-launch altitude 45,000 ft – 50,000 ft
Mothership B-52 Stratofortress or White Knight Two
Range, R 120 nm,
Empty Weight, WE 12,762.47 lbs
In-space maneuvering Reaction Control system (RCS) thrusters
Dock-undock location From behind and above X-69
Mechanism to dock-undock Nano-rack mechanism.
7.2. Appendix B: Proposed Aircraft Configuration:
Low wing
Land based aircraft
Conventional type with stabs and elevons
HS 130 airfoil for wing
Wing mounted empennage
Engine Type: Rocket motor engine
Engine Integration: Engine inside the fuselage from behind
Landing gear: Retractable gear
Nose-wheel landing gear
Rear gears attached to wings
7.3. Appendix C: Matlab code for take-off weight and comparison:
Following is the matlab code that can be used to estimate a precise take-off weight referring to
required comparison percentage weight less than 0.5%. The code helps to efficiently iterate for
WTO.
clc; clear all; close all;
%%
Wpl = 14000; % input your payload weight
A = -0.424; % regression coefficient, A for your aircraft
B = 1.161; % regression coefficient, B for your aircraft
fprintf('Wto w1 w2 w3 w4
w5 We comparison');
for Wto = 34000:100:35000 % create a for loop around the guessed take-off
weight
w1 = Wto*0.99;
%% Phase II - takeoff/Air launch, w2
w2 = w1*0.995;
%% Phase III - Climb, w3
44. 23
w3 = w2*0.9951;
%% Phase IV - Descent, w4 (glide)
w4 = w3*0.993;
%% Phase V - Landing, w5 (glide)
w5 = w4*0.995;
mff = w5/Wto;
Wf = (1-mff)*Wto;
r = ((log10(Wto))-A)/B;
We = 10^r;
Woe_tent = Wto-Wf-Wpl;
We_tent = Woe_tent-0.005*Wto-350;
comparison = (abs(We-We_tent)/((We_tent+We)/2))*100;
fprintf('n');
fprintf('%f %f %f %f %f %f %f %f
%f',Wto,w1,w2,w3,w4,w5,We,comparison);
end
Following piece of matlab code helps to solve a complicated equation 2.24 for takeoff weight
before beginning to calculate sensitivities:
clc; close all; clear all;
%%
A=-0.424;
B=1.161;
C=0.9635;
D=14350;
syms x
vpasolve(log10(x) == A+ B*log10(C*x-D),x)
Appendix D. AAA program initiation:
Step 1: Airplane configuration:
45. 24
Fig. 11: Airplane configuration
In this step, configuration of X-69 like wing, tail, spoiler configuration has been feed. For X-69, we will
have wing, 2 vertical tails at the rear end of the wing, horizontal tail attached to the wing.
Step 2: Propulsion:
Fig. 12: Propulsion
Propulsion system is specified in this step. X-69 will have one engine, buried in fuselage firing from behind
with an integral fuel tank and straight through inlet. Since rocketmotor engine used for X-69 is similar to
jet engine and there is no option for rocket engine so “Jet” is selected in “Propulsion” option.
46. 25
Step 3: Control surfaces:
Control surfaces for X-69 are similar to that of Spaceship Two. It will feathered wing configuration
with elevons, elevon trim tab on wing, differential stabilizers, also called as stabilators on horizontal
stabilizers and rudder with trim tab for vertical tail.
Fig. 13: Control surfaces-wing Fig. 14: Control surfaces – Horizontal Tail
47. 26
Fig. 15: Control surfaces - Vertical Tail
Step.4: Landing Gear:
As discussed in report 2, X-69 will have 3 landing gears. One attached to the fuselage near nose at
front and rest of the two will attached under wing with retraction capability.
49. Performance Constraints for X-69 CargoSat
AE 271 – Aircraft Design
Dr. Nikos Mourtos
Rushikesh Badgujar
San Jose State University
Charles W. Davidson College of Engineering
Aerospace Engineering
50. Contents
Introduction:..................................................................................................................................................4
Manual Calculation of Performance Constraints:.........................................................................................5
Climb Constraints: ....................................................................................................................................5
Drag Polar Estimation:..............................................................................................................................9
Landing Distance:...................................................................................................................................13
Selection of Propulsion System: .................................................................................................................14
Selection of the Propulsion System Type: ..............................................................................................14
Discussion:..................................................................................................................................................15
Parameters with major impact on design:...............................................................................................15
Conclusions and Recommendations: ..........................................................................................................15
References:..................................................................................................................................................16
Appendices:.................................................................................................................................................17
Appendix A: Density variation with altitude..........................................................................................17
Appendix B: Rate of climb computation using equation 2, 3 and 4. ......................................................18
Thrust to weight ratio and wing loading using equations 5, 6 and 7 ......................................................21
Appendix C: XFLR5 plots for lift, drag and glide ratios at high Mach numbers. ..................................24
51. List of Symbols:
F Force, N
M Mass, kg
a Acceleration, ft/sec2
or m/s2
T Thrust, N or lbf
W Gross Weight of X-69
D Drag force, N or lbf
γ Pullup or flight path angle, degree
t Time, seconds
RC Rate of climb at altitude h, ft/min
RC0 Rate of climb at sea level, ft/min
h Altitude, ft
habs Absolute ceiling altitude, ft
tcl Time of climb, seconds
V Resultant velocity, ft/sec
W/S Wing loading, psf
S Wing area, ft2
Swet Wetted area, ft2
A Aspect ratio
e Oswald efficiency
𝜌 Density of atmosphere, lbs/ft3
CD0 Coefficient of polar drag
L/Dmax Maximum glide ratio or lift/drag ratio
T/W Thrust to weight ratio
Re Reynolds number
VA Velocity of approach, ft/sec
VSL Stall speed at landing, ft/sec
SFL Field length, ft
CLmax Maximum coefficient of lift
(W/S)L Wing loading at landing, psf
CLmaxL Maximum coefficient of lift at landing
52. 1. Introduction:
This report investigates the performance constraints of X-69 based weight analysis and
previous flight data of similar airplanes. Performance of X-69 signifies parameters like rate of
climb, gliding and landing approach. This report also studies effect of different flight path angles
on rate of climb, T/W ratio and wing loading. Most of flight of X-69 depends on wing loading. In
climb phase wing loading is less prioritized since climb is solely performed by propulsion system
where wing configuration is feather locked to 0o
with no significant application in climb.
As discussed in previous reports, X-69 flight profile is divided into several stages as also
shown in figure below:
Fig. Flight profile of X-69 similar to that of VG’s Spaceship Two
a) Air launch and clean release: The mothership such as WhiteKnightTwo (WK2) makes a clean
release with pullup angle of 65o
.
b) Boost/Climb: Rocket engine fires up climbing X-69 to high altitudes. The total burn time is
expected to be approximately 90 seconds at which X-69 will attain 360,000 ft.
53. Rocket engine is the primary engine to boost X-69 to climb to 360,000 ft altitude at
supersonic speed. The boost phase of X-69 relies on Newton’s Second Law of Motion.
∑𝐹 = 𝑚𝑎 (1)
∑𝐹 is the summation of all external forces applied to the rocket, m is the mass of the X-69
accelerating with “a” ft/sec2
. The forces acting on X-69 during thrusting phase (climbing) of flight
are its weight (W), Thrust (T) and aerodynamic drag (D). The effectiveness of thrust varies as the
vertical component propels the vehicle to the target altitude and depends on flight path (pullup)
angle with weight continuously changing due to the burning of rocket fuel.
c) Coast: After 90 seconds of boost and reaching apogee, X-69 performs the desired mission to
deploy satellites using precise maneuverability with RCS thrusters.
d) Re-entry: Using RCS thrusters to orient its attitude for re-entry. The reentry phase is up to
80,000 ft – 85,000 ft. Reentry is accompanied by changing wing configuration to feathered
state where wing feather gets locked to 60o
using pneumatic system.
e) Descend and Glide: X-69 decelerates using aerodynamic drag with efficient gliding
performance. The wing is designed to generate more and stable drag to kill the reentry speeds.
f) Approach and Landing: X-69 makes an approach for landing at subsonic speed. Further
landing performance is studied in performance constraints section. For this analysis, Mojave
Airspace and Spaceport is considered to simplify and compare the analysis with spaceship two
since flights of spaceship two were performed on this airport.
2. Manual Calculation of Performance Constraints:
Manual calculations for X-69 performance constraints are performed referring to MIL-C-
005011B. Since X-69 has potential for high maneuverability and climb rate at supersonic speed,
its climbing characteristics are considered similar to that of fighter planes with steep flight path
angles, γ.
2.1. Climb Constraints:
Considering total climb time during boost phase, t = 90 sec referring to previous similar
airplanes and their flight profiles. From Roskam book, Part I, section 3.4.10 describes the equation
for rate of climb at certain altitude. The equation is as follows:
𝑅𝐶 = 𝑅𝐶0(1 −
ℎ
ℎ 𝑎𝑏𝑠
) (2)
Where, RC = rate of climb at altitude, h in fpm
RC0 = rate of climb at sea level in fpm
Since we don’t know rate of climb at sea level, it can be calculated as:
𝑅𝐶0 = (
ℎ 𝑎𝑏𝑠
𝑡 𝑐𝑙
) . ln(1 −
ℎ
ℎ 𝑎𝑏𝑠
)−1
(3)
Absolute ceiling, habs for X-69 is 360,000 ft
Since X-69 attains a steep flight angle, γ = 65o
, 75o
, 80o
, 85o
and 90o
, equation 3.37 from
Roskam book, can be used to calculate V.
54. 𝑅𝐶 = 𝑉𝑠𝑖𝑛𝛾 (4)
Using excel, rate of climb and respective velocities is computed with rise in altitude and with respect to
time using equations 2, 3 and 4. This computed data is tabulated in Appendix B in table. B.1.
Fig. 2 shows location of X-69 at certain altitude at specific time in 90 seconds of climb profile.
Fig.1. Altitude v/s rate of climb
Fig. 2. Altitude of X-69 v/s time
If the climb rate is to be maximized, L/D needs to be maximized.
0
50000
100000
150000
200000
250000
300000
350000
400000
0 20000 40000 60000 80000 100000
altitude,ft
Rate of climb, ft/min
altitude v/s rate of climb
0
50000
100000
150000
200000
250000
300000
350000
400000
0 20 40 60 80 100
Altitude,ft
time, sec
Altitude v/s time
55. Hence using equations 3.34, 3.35 and 3.36 plot between thrust to weight ratio, T/W and wing loading can
be computed.
𝑉 =√
2×
𝑊
𝑆
𝜌√ 𝜋𝐴𝑒.𝐶 𝐷 𝑜
(5)
𝐶 𝐷 𝑜
= (
𝜋𝐴𝑒
2×((𝐿/𝐷) 𝑚𝑎𝑥)
)2
(6)
A (aspect ratio of X-69, is considered same as that of spaceship two) = 1.62
e (Oswald efficiency for supersonic flight regime) = 0.3-0.5
and referring to the flight log of spaceship two, (L/D)max = 7
Table A.1 in Appendix A documents density of air at higher altitudes. Density almost goes to zero at and
above 150,000 ft altitude where X-69 will experience zero drag.
Therefore,
𝐶 𝐷 𝑜
= (
𝜋 × 1.62 × 0.5
2 × (7)
)
2
𝐶 𝐷 𝑜
= 0.01298
Hence substituting CD0 and V at increasing altitude and for flight path angle, we get respective wing
loading that can be referred from Appendix B, in table. B.2.
Using equation 3.34 from Roskam book, we can calculate T/W for maximum L/D as follows:
𝑅𝐶 = 𝑉{
𝑇
𝑊
− (
1
𝐿
𝐷
)}
𝑇
𝑊
=
𝑅𝐶
𝑉
+
1
𝐿
𝐷
(7)
This thrust to weight ratio refers to respective wing loading which is eventually outcome of respective
velocities and climb rates considering while calculating through above equations.
Fig. 1 is the plot of result of equation (5) and (7) describing T/W for different wing loadings with
respect to flight path angles. As we can observe from the plot that T/W reduces with increase in wing
loading. Although comparatively T/W is higher for lower pullup angles due to obvious reason. When pullup
angle increases, vertical component of weight increases resulting into less T/W ratio but following similar
patterns with respect to wing loading. Wing loading is zero above 150,000 ft altitude can be referred from
Appendix B, table B.2.
57. Fig. 5. Altitude v/s wing loading.
Fig. 4 describes effect of wing loading on velocity. With higher wing loading, velocity
drops significantly with similar pattern for different flight path angle. Similarly, fig. 5 describes
how wing loading is affected at high altitude. In both the cases, velocity is zero at zero wing loading
since this case is when X-69 is at high altitude where there is no atmosphere for drag. Similarly, at
high altitudes wing loading is zero and rises as the altitude drops and earth’s atmosphere engulfs
X-69.
Eventually it is less necessary to worry about wing loading at climb phase since wing has
no function in climbing. Climb is solely performed by primary propulsion system providing
continuous thrust.
2.2. Drag Polar Estimation:
Airfoil HS130 is considered for wing and feathers. HS130 is well-known for its gliding
performance at high-altitudes. While entering to glide phase, wing feather gets locked at 60o
. This
configuration provides ample amount of drag to decelerate while descending. The feather is locked
into the place by a set latches that is driven by pneumatic pistons as shown in fig. 6.
50000
70000
90000
110000
130000
150000
170000
190000
0 50 100 150 200 250 300 350 400
Altitude,ft
Wing Loading, W/S, lb/ft2
Altitude v/s wing loading
altitude v/s W/S @ 65deg
altitude v/s W/S @ 75deg
altitude v/s W/S @ 80deg
altitude v/s W/S @ 85deg
altitude v/s W/S @ 90deg
58. Fig. 6. Feather retraction and deployment
XFLR5 have been used to estimate the lift, drag and glide ratio at various angle of attacks.
This analysis is performed on HS130 airfoil to examine its aerodynamic performance. Fig. 7 shows
the imported airfoil and pressure distribution at 0o
angle of attack at Mach 0 and Reynolds number
100,000.
Fig. 7. Pressure distribution on HS130 Airfoil
Drag polar for climbing is estimated in climbing section. We get CD0 = 0.013 from equation
6. Using XFLR5 lift and drag distribution at rising angle of attacks at various Reynolds number is
analyzed. In this case Ncrit is considered as standard of e9
.
59. Fig. 8. Inputs for batch foil analysis
Fig. 9. Coefficient of lift v/s angle of attack at various Reynolds numbers and Mach 0.0
60. Fig. 10. Lift v/s drag coefficient at similar conditions as fig. 9
Fig. 11. Cl/Cd (glide ratio) v/s angle of attack at similar conditions as fig. 9
61. Fig. 9 gives the estimate of Clmax which lies between 1.0 to 1.2 at rising Reynolds number
at approximately 12o
to 12.5o
angle of attack. This analysis resembles when X-69 descends from
80,000 ft of altitude.
Plots for Mach 0.2 and 0.3 can referred from Appendix C. X-69 is expected decelerate from
reentry to approach velocity VA by the time it reaches to landing stage. This approach speed should
be between 130 knots to 140 knots or 0.2 to 0.3 Mach. Hence using XFLR5 in appendix C, plots
for lift and drag coefficient and angle of attack are computed which shows wing configuration still
has stable aerodynamic parameters at similar Reynolds number and different speeds.
2.3. Landing Distance:
Landing distance sizing for X-69 is performed with respect to FAR 25 regulations. From
the previous flight logs of Spaceship Two and its location for flight tests, Mojave Air and Space
Port is considered for landing. Hence parameters related field length will be considered with
respect to runways of this air base. It is required to size X-69 for a landing field length of
averagely estimated 7000 ft at sea level on a standard day.
It may be assumed that WL = 0.85WTO.
From Roskam, equation 3.16
𝑆 𝐹𝐿 = 0.3𝑉𝐴
2
(8)
Hence,
𝑉𝐴 = √(
7000
0.3
)
𝑉𝐴 = 152.7𝑘𝑡𝑠
With equation 3.17 from Roskam, which is as follows
𝑉𝑆 𝐿
=
𝑉 𝐴
1.2
(9)
𝑉𝑆 𝐿
=
152.753
1.2
𝑉𝑆 𝐿
= 127.3𝑘𝑡𝑠
Using equation 3.1 from Roskam,
𝑉𝑆 =√
2𝑊/𝑆
𝜌𝐶 𝐿𝑚𝑎𝑥
(10)
Therefore, substituting VSL stall speed for landing in above equation in ft/sec, we get
relationship between CLmax and W/S.
(
𝑊
𝑆
)
𝐿
= (127.3 × 1.688
𝑓𝑡
sec
)
2
× 0.002378 ×
𝐶𝐿 𝑚𝑎𝑥 𝐿
2
62. (
𝑊
𝑆
)
𝐿
= 54.9𝐶𝐿 𝑚𝑎𝑥 𝐿
(11)
From table 3.1 of Roskam, CLmaxL is considered referring to Supersonic Cruise Airplanes which
is in the range of 1.8-2.2.
Hence table. 1 shows wing loading at given CLmaxL using equation (11). Wing loadings below
for varying CLmaxL are when X-69 approaches for landing.
Table. 1. Wing loading for given 𝐶𝐿 𝑚𝑎𝑥 𝐿
𝐶𝐿 𝑚𝑎𝑥 𝐿
(W/S)L
1.8 98.8 psf
2 109.8 psf
2.2 120.8 psf
3. Selection of Propulsion System:
3.1. Selection of the Propulsion System Type:
Selection of propulsion system type is based on various factors that are described and specified
for Rocket Motor Two (RM2), the solid-liquid hybrid rocket engine that will be used for X-69 as
follows:
Maximum speed: Desired cruise speed or maximum speed comes into play for engine type
selection. According to mission specifications of X-69, it will climb at Mach 3.0 – 3.25.
Maximum operating Altitude: After clean release from mothership at 65o
pull-up angle to
50,000 ft of altitude, X-69 will boost to climb to altitude of 360,000 ft in approximately 90 sec.
Although air density at 50,000 ft and above is less, X-69 will still experience considerable
amount of drag due atmosphere and gravity. To overcome this drag, it is necessary to use the
engine with high thrust.
Range and Economy: Through the flight course of X-69, engine will be used only for climb.
Coasting and re-entry will be performed by reaction control thruster system which is considered
part of payload and wing as it descends. Hence technically speaking, range of X-69 will only
be one-way trip to the coasting altitude. At this point range requirement is not significant.
Hence a solid-liquid hybrid rocket engine will be the primary propulsion system for X-69.
Rocket Motor Two uses nitrous oxide N2O as oxidizer and hydroxyl-terminated polybutadiene
(HTPB) as a solid propellant, the combination used in Spaceship One and Spaceship Two.
Although to increase the efficiency, manufacturers of RM2 are pursuing to study other solid
propellants such as thermoplastic polyamide (nylon).
Propulsion system and specifications:
Primary propulsion system: RocketMotorTwo manufactured by SNC generates 60,000 lbf of
thrust and 250 sec of specific impulse. The total burn time of RM2 for X-69 is 90 seconds at
which X-69 attains 360,000 ft of altitude.
63. 4. Discussion:
X-69 is designed to climb at supersonic speed to high altitudes and coast near LEO to deploy
cubesats. In report 3 we estimated takeoff weight of X-69 of 36505 lbs. To estimate precise climb, glide
and landing, calculations for performance constraints have been performed manually. From fig. 3, we
can see that thrust to weight ratio varies in similar pattern with respect to wing loading at various
possible pullup angles. In previous flights spaceship one and spaceship two, standard pullup angle at
clean release from mothership and thereon boost is 65o
to 75o
. Also we can see from plots of both
velocity and T/W in fig. 3 and 4 respectively, with low T/W, X-69 can still achieve high velocities
increasing the rate of climb with almost same wing loading. Although while descending, wing
configuration and high drag helps to decelerate and hence stabilize X-69 before approach to land.
Parameters with major impact on design:
Rate of climb is one of the significant parameters that has major impact on design and hence the
performance. Consequently, it is also important to have optimal primary propulsion system that can
provide continuous thrust to complete the boost phase of about 90 seconds. Advantage of using hybrid
rocket propulsion is that thrust can be controlled by pilot maintaining the oxidizer flow during
combustion unlike solid propellant rockets. In this analysis it is assumed that the secondary propulsion
system that is required for orbital maneuvering is optimal and will only come into play at coasting and
re-entry initiating phase.
5. Conclusions and Recommendations:
Theoretically and based calculations and previous flights of similar airplanes, it is desired to keep
the flight path angle or pullup angle of 65o
. This reduces wing loading as compared to higher pullup
angles. For estimation of drag polar, L/Dmax of 7:1 is considered referring to flight log of spaceship two
assuming same aspect ratio of 1.62. It is desired to perform wind tunnel analysis on aerodynamic
performance of wing configuration to seek for more simple configuration expecting controllable drag
while descending. Also since this study limits to a certain landing zone, it is crucial to perform landing
constraints analysis based on other parameters related to landing including temperature, density of
atmosphere, elevation of runway from sea level, etc.
64. 6. References:
(n.d.). Retrieved from
http://www.petervis.com/interests/published/Spaceshiptwo/Spaceshiptwo_Rocket.html.
(n.d.). Retrieved from http://www.nbcnews.com/storyline/virgin-voyage/how-spaceshiptwos-
feathered-wings-were-supposed-work-n240256.
(n.d.). Retrieved from http://bagera3005.deviantart.com/art/White-Knight-SpaceShip-One-
158659309.
(n.d.). Retrieved from https://en.wikipedia.org/wiki/SpaceShipOne.
Airfoil Database. (n.d.). Retrieved from https://www.aerodesign.de/english/profile/profile_s.htm.
Airfoils and Airflow. (n.d.). Retrieved from https://www.av8n.com/how/htm/airfoils.html.
Donald Greer, hamory, P., Krake, K., & Drela, M. (n.d.). Design and Predictions for a High-
Altitude (Low-Reynolds-Number) Aerodynamic Flight Experiment.
Evans, M. (2013). The X-15 Rocket Planes. In M. Evans, The X-15 Rocket Planes.
FAA. (n.d.). Aerodynamics of Flight. In Gliding Flight Handbook.
Flight Training Center. (n.d.). Retrieved from http://flighttrainingcenters.com/training-aids/multi-
engine/engine-out-procedures/.
Global Aircraft. (n.d.). Retrieved from https://www.globalaircraft.org/planes/x-15_hyper.pl.
https://en.wikipedia.org/wiki/North_American_X-15. (n.d.). Retrieved from
https://en.wikipedia.org/wiki/North_American_X-15.
NASA. (n.d.). A NASA Guide to engines.
NASA factsheet. (n.d.). Retrieved from
https://www.nasa.gov/centers/armstrong/news/FactSheets/FS-052-DFRC.html.
Scaled Composites. (n.d.). Retrieved from http://www.scaled.com/projects/tierone/.
Space Flight Laboratory. (n.d.). Retrieved from http://utias-sfl.net/?page_id=87.
Space.com. (n.d.). Retrieved from http://www.space.com/30245-x37b-military-space-plane-100-
days.html.
Virgin Galactic. (n.d.). Retrieved from http://www.virgingalactic.com/human-spaceflight/our-
vehicles/.
Virgin galactic fact sheet. (n.d.). Retrieved from
http://www.galacticexperiencesbydeprez.com/pdf/vg_vehicles_fact_sheet101411.pdf.
wired.com. (n.d.). Retrieved from https://www.wired.com/2010/10/test-pilot-describes-first-glide-
flight-of-spaceshiptwo/.
65. 7. Appendices:
7.1 Appendix A: Density variation with altitude
Density of air almost goes to zero at and above 150,000 ft altitude.
Table. A.1: Density at altitudes
h D, lb/ft3
48000 0.013909768
50000 0.012734849
55000 0.010128612
60000 0.007951502
65000 0.006151895
70000 0.004681823
75000 0.003496919
80000 0.002556358
85000 0.001822797
90000 0.001262307
95000 0.000844305
100000 0.000541485
105000 0.000329733
110000 0.00018805
115000 9.84539E-05
120000 4.58831E-05
125000 1.80836E-05
130000 5.48162E-06
135000 1.0372E-06
140000 6.47948E-08
145000 1.48867E-12
150000 0
155000 0
160000 0
72. 7.4 Appendix C: XFLR5 plots for lift, drag and glide ratios at high Mach numbers.
Fig.C.1: Lift distribution at increasing angle of attacks at Mach 0.2
Fig. C.2: Lift v/s drag at Mach 0.2
73. Fig. C.3: Glide ratio at Mach 0.2 at increasing angle attack
Fig. C.4: Lift distribution at increasing AoA at Mach 0.3
74. Fig. C.5: lift v/s drag at Mach 0.3
Fig. C.6 Glide ratio at Mach 0.3
75. Fuselage design for X-69 CargoSat
AE 271 – Aircraft Design
Dr. Nikos Mourtos
Rushikesh Badgujar
San Jose State University
Charles W. Davidson College of Engineering
Aerospace Engineering
76. Contents
1. Introduction: .......................................................................................................................................3
2. Layout design of the cockpit:.............................................................................................................3
2.1. 2D CAD models...........................................................................................................................4
2.2. 3D CAD models:..........................................................................................................................5
3. Layout design of the fuselage:............................................................................................................5
3.1. 2D CAD models................................................................................................................................6
3.2. 3D CAD models:...............................................................................................................................7
4. Discussion: ...........................................................................................................................................8
5. References:.........................................................................................................................................10
77. 1. Introduction:
Before starting layout design of cockpit and fuselage, it is important to revise the mission
specification based on which a comprehensive overall configuration can be interpreted.
Table.1: Mission Specification
Crew 2 pilots
Weight of crew 350 lbs, 175 lbs each
Payload Cubesats and small satellites
Payload weight 1,300 lbs
Deployment altitude 360,000 ft to 400,000 ft
Wing configuration Low wing
Deployment direction from fuselage Upwards with retractable door mechanism
In-fuselage major components NanoRack mechanism to swiftly deploy satellites
Cockpit and fuselage avionics
The strategy of low wing is for convenient deployment of satellites from above. Wing and nose
of X-69 have RCS thrusters to control roll and pitch and yaw respectively. Considering this project
as a prototype, at this time X-69 can contain 24 cubesats with various categories. Before studying
cockpit and fuselage layout, it is important to understand type of payload and their mission
requirements. Upon research it is found that some cubesats require spacecraft assist deployment
whereas some have self-deployment mechanisms either rails or small solid propellant rockets.
Development of cubesats started with 1U and 3U sizes build by various resources. After
successful services delivered by this concept, demand for bigger sized cubesats increased due to
which sizes of cubesats now ranges from 1U (10cm x 10cm x 10cm) to 27U (34cm x 35cm x 36cm).
2. Layout design of the cockpit:
Initial proposed X-69 will have cockpit with conventional layout. As shown in fig. 4. 2 pilot
seats approximately aligned at 15inches from the center line running through the nose. Nose is
essentially long to house RCS thrusters for pitch and yaw control while coasting in space. Fuel
management system for RCS thrusters for both at nose and wing is provided through the layer
under dashboard layer. Blank space behind the crew in cockpit is to accommodate accessories such
as washrooms, facility to store beverages, etc. Unlike conventional perspective to load and unload
crew into the cockpit, it is quite feasible to use seat elevators that will reach down to the ground for
crew to get off of the cockpit.
78. 2.1.2D CAD models:
Fig. 1. Top view of cockpit
Fig. 2. Front view of cockpit
Fig. 3. Side view of cockpit
Fig. 4. Rear view of cockpit with seat.
Fig. 5. 2D sketch of Cockpit layout