Minimum impulse guidance Interim scientific report
Minimum fuel guidance for time-open rendezvous, orbit transfer, and orbit transfer with tangential nominal impulses
Edelbaum, T. N.↗
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Optimization of a solar-electric-propulsion planetary orbiter spacecraft.
Solar electric propulsion spacecraft design for 1971 Mars orbital rendezvous, optimizing net spacecraft mass with respect to several constraints
Sauer, C. G., Jr.↗
A study of an Extended Lunar Orbital Rendezovous /ELOR/ mission. Volume 1 - Technical analysis Final report
Feasibility of Extended Lunar Orbital Rendezvous /ELOR/ for use in lunar application programs
Carpenter, R. B., Jr.↗
Apollo experience report: The docking system
The decision to accomplish the lunar landing mission by use of the lunar orbit rendezvous technique required that a docking system be developed to allow: (1) spacecraft modules to be structurally joined, (2) intravehicular transfer of the crew and equipment, and (3) separation of the modules. The basic design criteria of the docking system, the evolution process, and the various docking concepts considered for the Apollo program are presented. Docking systems that were considered for the Apollo program included both impact and nonimpact systems; a probe and drogue impact system was selected. Physical and functional descriptions of the probe and drogue, the crew transfer tunnel, and docking ring latches are presented for both the early configuration and the present configuration as influenced by the development and qualification test programs. In addition, preflight checkout activity and mission performance of the system are discussed.
Langley, R. D.↗
Guidance and control of flight vehicles
Progress reports on guidance and attitude control mechanisms of different flight vehicles are presented. The vehicles considered include orbiting spacecraft, supersonic aircraft, and general aviation aircraft. Data also cover orbital transfer using low thrust, automatic landing logic for aircraft, optimal and three dimensional turns for supersonic aircraft, and orbital rendezvous.
Breakwell, J. V.↗
Development and applications of an orbital insertion surface for the space shuttle orbiter/tug
The concept of a space shuttle orbit/tug orbital insertion surface is developed. Practical applications in the area of ascent targeting, on-orbit rendezvous targeting, payloads/OMS propellant off-loading, and mission analysis are identified.
Deaton, A. W.↗
Apollo experience report: Lunar module structural subsystem
In the Apollo Program, the lunar-orbit rendezvous mode was used for the lunar-landing mission. To accomplish the lunar landing, a lunar module spacecraft was built. A description of the design requirements for the structural subsystem and of the structural configuration and the method of design verification are given. A discussion is presented of several problems encountered and the corrective actions taken during the designing, manufacturing, and testing of the lunar module.
Weiss, S. P.↗
Astronomy sortie missions definition study. Volume 3, book 2: Appendix: Design analysis and trade studies
Backup or supporting data for the design analyses and trade studies which defined the astronomy sortie missions are presented. The subjects discussed are: (1) configuration of space shuttle orbiter, (2) electronic subsystems, (3) electric power requirements, and (4) payload requirements. Mathematical models are developed to illustrate the orbital rendezvous capabilities.
Source record↗
Use of Green's functions in the numerical solution of two-point boundary value problems
This study investigates the use of Green's functions in the numerical solution of the two-point boundary value problem. The first part deals with the role of the Green's function in solving both linear and nonlinear second order ordinary differential equations with boundary conditions and systems of such equations. The second part describes procedures for numerical construction of Green's functions and considers briefly the conditions for their existence. Finally, there is a description of some numerical experiments using nonlinear problems for which the known existence, uniqueness or convergence theorems do not apply. Examples here include some problems in finding rendezvous orbits of the restricted three body system.
Gallaher, L. J.↗
Apollo experience report guidance and control systems
The Apollo guidance and control systems for both the command module and the lunar module are described in a summary report. General functional requirements are discussed, and general functional descriptions of the various subsystems and their interfaces are provided. The differences between the original in-flight maintenance concept and the final lunar-orbital-rendezvous concept are discussed, and the background in philosophy, the system development, and the reasons for the change in concept are chronologically presented. Block diagrams showing the command module guidance and control system under each concept are included. Significant conclusions and recommendations contained in more detailed reports on specific areas of the guidance and control systems are included.
Wilson, R. E., Jr.↗
Nominal profile refinements report: Targets in 150 and 190 nautical mile circular orbits
Refinements for 150 and 190 n.mi. circular target orbits were examined along with the potential advantages of ground-hold as a means of reducing the phasing times associated with low orbit rendezvous.
Kreiter, T. J.↗
Minimum-fuel rocket trajectories involving intermediate-thrust arcs
The optimal trajectories in the neighborhood of an optimal intermediate-thrust arc are investigated for the minimum-fuel orbit rendezvous problem with fixed specific impulse. Since such an arc is singular, the thrust acceleration magnitude being the singular control component, a second-variation analysis leads to the identification of a field of neighboring, singular arcs in a state space of dimension four rather than six, provided that a suitable Jacobi condition is met. A given neighboring initial six-dimensional state vector does not generally lie on a neighboring singular arc, and junction onto the appropriate singular arc must be accomplished by a short period of strong variations in the acceleration. The neighboring singular arc meets the final condition in 4 dimensions, rather than 6 dimensions, and rendezvous must be completed by another, terminal short period of strong variations in the acceleration. Implications for midcourse guidance near a singular arc are discussed.
Breakwell, J. V.↗
Use of radio equipment for Space Shuttle navigation
Space Shuttle navigation is defined, in a narrow sense, as the task of maintaining adequate knowledge of vehicle position and velocity. The state of the Orbiter in terms of this goal is described by a vector of at least six elements, three for position and three for velocity, at a given time. These are referred to as the 'state vector' or simply as the state. Shuttle navigation will relay on a blend of ground-based and onboard systems. The onboard systems will be capable of state propagation at all times and will perform state determination during the latter part of the entry from orbit. The ground-based system will be capable of accurate state propagation for free-flight phases. Radio communication will be necessary in order for the ground and onboard capabilities to work as a coordinated system. The use of radio equipment for state determination is discussed for the ascent, orbit, rendezvous, descent, and abort phases of Shuttle missions.
Schiesser, E. R.↗
Orbit lifetime characteristics for Space Station
The factors that influence the orbital lifetime characteristics of the NASA Space Station are discussed. These include altitude, attitude, launch date, ballistic coefficient, and the presence of large articulating solar arrays. Examples from previous program systems studies are presented that illustrate how each factor affects Station orbit lifetime. The effect of atmospheric density models on orbit lifetime predictions is addressed along with the uncertainty of these predictions using current trajectory analysis of the Long Duration Exposure Facility spacecraft. Finally, nominal reboost altitude profiles and fuel requirement considerations are presented for implementing a reboost strategy based on planned Shuttle Orbiter rendezvous strategy and contingency considerations.
Deryder, L.↗
Mars direct - A coherent architecture for the Space Exploration Initiative
The coherency of the conventional earth-orbital assembly/Mars orbital rendezvous is analyzed in terms of the Space Exploration Initiative (SEI) architecture and an alternate approach is given. The coherency - defined as a combination of simplicity, robustness, and cost-effectiveness - of the conventional SEI is found to be inadequate. The Mars Direct plan is described in detail and proposed as an efficient alternative for both Martian and lunar applications. In situ propellant production on the target-planet's surface is described, and candidate vehicle designs are set forth. Objectives and problems associated with a manned mission are outlined, and details regarding surface activities and support are given. Nuclear thermal rocket propulsion is determined to be an efficient way to improve mission capability, and the Mars Direct SEI architecture can integrate such technology. The Mars Direct SEI is concluded to be a more effective plan than the conventional rendezvous mission.
Zubrin, Robert M.↗
A space transfer vehicle for lunar missions
Two 'low profile' strategies for supporting the lunar outpost concept developed by NASA-JSC Planetary Surface Systems are examined. The first strategy involves the design of missions that require only a single launch of a derivative 'National Launch System' (NLS) vehicle. The launch capacity of the derivatives range from 80 to 140 percent of the Apollo/Saturn capacity and use only conventional cryogenic or stable chemical propulsion. The second strategy involves the design of missions that only require the currently planned NASA NLS vehicle to be built. This strategy uses either one or two launches with an earth and/or lunar orbit rendezvous for stage assembly. Advance propulsion concepts, in addition to cryogenics, for the translunar injection stage are addressed.
Andrews, Dana↗
Project Columbiad: Mission to the Moon. Book 2, volume 3: Stage configuration designs; volume 4: Program plan
The Earth Orbital Rendezvous (EOR) configuration for the piloted mission is composed of three propulsive elements in addition to the Crew Module (CM): Primary Trans-Lunar Injection (PTLI), Lunar Braking Module (LBM), and Earth Return Module (ERM). The precursor mission is also composed of three propulsive elements in addition to its surface payloads: PTLI, LBM and the Payload Landing Module (PLM). Refer to Volume 1, Section 5.1 and 5.2 for a break-up of the different stages into the four launches. A quick summary is as follows: PTLI is on Launch 1 and 3 while the LBM, PLM, and surface payloads are on Launch 2 and another LBM, ERM, and CM on Launch 4. The precursor mission is designed to be as modular as possible with the piloted mission for developmental cost considerations. The following topics are discussed: launch vehicle description; primary trans-lunar injection stage; lunar braking module; earth return module; crew module; payload landing module; and surface payload description.
Source record↗
Project Columbiad: Mission to the Moon. Book 1: Executive Summary. Volume 1: Mission trade studies and requirements. Volume 2: Subsystem trade studies and selection
In response to the Report of the Advisory Committee on the future of the U.S. Space Program and a request from NASA's Exploration Office, the MIT Hunsaker Aerospace Corporation (HAC) conducted a feasibility study, known as Project Columbiad, on reestablishing human presence on the Moon before the year 2000. The mission criteria established were to transport a four person crew to the lunar surface at any latitude and back to Earth with a 14-28 day stay on the lunar surface. Safety followed by cost of the Columbiad Mission were the top level priorities of HAC. The resulting design has a precursor mission that emplaces the required surface payloads before the piloted mission arrives. Both the precursor and piloted missions require two National Launch System (NLS) launches. Both the precursor and piloted mission have an Earth orbit rendezvous (EOR) with a direct transit to the Moon post-EOR. The piloted mission returns to Earth via a direct transit. Included among the surface payloads preemplaced are a habitat, solar power plant (including fuel cells for the lunar night), lunar rover, and mechanisms used to cover the habitat with regolith (lunar soil) in order to protect the crew members from severe solar flare radiation.
Clarke, Michael↗