Kistler had been the farthest along and the most technically feasible of the privately-funded commercial launch vehicle projects of the late 1990's. Kistler followed the 'lean company' model. The structure and systems of the launch vehicle were subcontracted to major US aerospace companies. The vehicle itself used surplus Russian engines developed for the never-flown N1F launch vehicle. The first stage of the vehicle would fly back to the launch site; the second would orbit Earth before returning. Recovery of the stages would be by parachutes and air bag. Launches from both the Nevada Nuclear Test Range and the Australia Woomera launch sites were considered. Interestingly, this was the most economic scheme for a recoverable launch vehicle as identified by Mishin, chief designer of the N1. The company had raised $440 million and development was well underway when the Asian financial crisis and Iridium bankruptcy scared away investors.
The Kistler K-1 was a two-stage vehicle projected for full reusability of both stages. It was 36.9m in overall length, 6.7m in diameter and weighed 382,300 kg at lift-off. The vehicle was powered by Aerojet/Russian NK-33 and NK-43 engines and was designed to be reused 100 times. Maximum payload was 4,500 kg to a 200 km, 45 degree inclination orbit. Typical MEO payload was 2,000 kg to a 900 km, 60 degree orbit.
The Kistler Aerospace development schedule consisted of three phases. Each phase included the completion of specific financing goals and the achievement of planned technology milestones. Kistler was well into Phase II before experiencing funding problems.
Phase I: Preliminary Design - 1994-1995
By the end of 1995, Kistler Aerospace had completed Phase I of its development plan including the following activities.
Hired key members of the K-1 technical team. Completed the K-1 preliminary baseline design in September 1995. Raised additional capital from private sources. Engaged in negotiations with potential investors, including potential strategic partners, and potential contractors. Completed an assessment of the global satellite launch market and created a strategic plan for the K-1's entry into the global satellite launch business. Completed a series of presentations to K-1 customer prospects in the U.S., Europe, and Asia.
Phase II: Development and Testing - 1996-1999
Completed the K-1 vehicle development and ground facilities work packages and signed preliminary contracts with several major systems and sub-systems contractors. Completed a series of critical design reviews with K-1 contractors. Raised additional private financing. Signed a Memorandum of Understanding with the State of Nevada to establish a United States launch site. Received 46 Russian NK-33 and NK-43 engines from Aerojet. Obtained a Right of First Refusal for all remaining Russian NK-33 and NK-43 engines.
Space Systems/Loral awarded a contract for ten launches with a value in excess of $100 million.
Executed contracts with major team members:
As funding becomes available, Kistler planned to complete Phase II. This would involve completing the detailed design, fabrication, assembly, integration, and ground testing of the K-1 aerospace vehicles. The flight test program was designed to demonstrate the capabilities and performance of K-1 vehicles, including guidance and control systems, ground and flight hardware, ground and flight operations procedures, thermal protection systems, and parachute and airbag landing systems.
Phase III: Commercial Operations - Beyond 2010
Routine K-1 fleet orbital operations, delivering customer payloads to low earth orbit, was originally planned to begin in 2000. This date continues to slip based on the availability of funding to complete development.
K-1 Technical Description
The K-1 vehicle was a two-stage, fully reusable aerospace vehicle. The overall K-1 vehicle was 36.9 m long and weighed 382,300 kg at lift-off.
The first stage, or Launch Assist Platform (LAP), was 18.3 m long, 6.7 m in diameter, and weighed 250,500 kg at launch. The second stage, or Orbital Vehicle (OV), was 18.6 m long, had a cylindrical diameter of 4.3 m, and weighed 131,800 kg fully-fuelled. Each stage carried its own suite of redundant avionics and operated autonomously.
The three liquid oxygen (LOX)/ kerosene engines of the LAP were high performance, robust engines which combined the best of Russian and U.S. technologies. The two AJ26-58 engines and one AJ26-59 engine included the fully developed, extensively tested core of the NK33/43 engines originally built for the Russian Manned Moon Program. The engines were modified to include modern U.S. electronic controllers, ignition systems, control valves, and thrust vector control (TVC) systems.
The three LAP engines provide 4,540 kN of total thrust at lift-off. These engines have an expansion ratio of 27:1 and are capable of being hydraulically gimballed to ± 6-degrees in pitch and yaw. Stage separation occurred at 139 seconds into the mission at an altitude of 43.2 km and a velocity of 1.22 km/sec. The LAP centre engine re-ignited 4.4 seconds after separation to return the stage to the launch site.
Other major subassemblies of the LAP include the composite interstage, aluminium liquid oxygen (LOX) tank, composite intertank, aluminium fuel tank, a retention tank to contain the LOX necessary for the flyback to the launch site, and the composite aft skirt. Four cold gas thruster pods provided attitude control during LAP separated flight. Tank pressurisation was performed using helium. The avionics system was located in the intertank compartment.
Two drogue parachutes and two clusters of three main parachutes were used to decelerate the LAP for a soft touchdown using four low-pressure airbags.
The OV used one AJ26-60 engine for main propulsion. The AJ26-60 provided 1,760 kN vacuum thrust with an expansion ratio of 80:1. The engine was hydraulically gimballed to provide thrust vector control. 7.3 seconds after LAP separation, the AJ26-60 engine ignited for a typical 230 second burn to place the vehicle in an elliptical orbit with an apogee at the deployment altitude. OV main engine cut-off typically would occur at an altitude of 94.4 km and a velocity of 7.8 km/sec. Following a coast to apogee, the LOX/ethanol Orbital Manoeuvring System (OMS) fired to circularise the orbit.
Following OV orbit insertion, the attitude control system (ACS) was used to damp out any residual attitude rates and orient the vehicle in the required attitude. The payload module dome was opened and payload deployment was initiated. Separation data were immediately sent to the ground operations centre using TDRSS.
OV post-deployment manoeuvres were designed to eliminate the possibility of re-contact with or contamination of the customer's payload. After payload deployment, the OMS fired again to place the OV into a phasing orbit with the correct period for re-entry. Following a second coast phase of up to 22 hours, the vehicle reoriented and performed de-orbit burn with the OMS, and re-entered the earth's atmosphere. The OV flew a guided re-entry trajectory to the launch site. A high-altitude stabilisation chute was deployed at Mach 2.5, followed by deployment of a single drogue and three main chutes. The main parachutes decelerated the stage for a soft touchdown using four low-pressure airbags.
Other major subassemblies of the OV included the composite forward skirt, aluminium LOX tank, composite intertank, composite fuel tank and composite flare. Tank pressurisation was performed using helium. Four cold gas thruster pods provide stage attitude control during separated flight. The avionics system was located in the forward compartment behind the payload module.
The K-1 vehicle offered the customer two payload module configurations: the Standard Payload Module (SPM) and the Extended Payload Module (EPM). The payload modules were fabricated from composite materials and used redundant, high-reliability mechanisms. Both payload module configurations incorporated interior acoustic absorption blankets and had integral pre-launch environmental control systems.
The payload modules were vertically integrated with the payload in the Payload Processing Facility. Payload modules were interchangeable to provide maximum flight schedule flexibility.
Each stage of the K-1 vehicle was completely autonomous from launch to landing. Stage guidance and control were provided by a triple-redundant, fault-tolerant avionics architecture.
Vehicle position, velocity and orientation data were provided by three integrated Global Positioning System (GPS) / Inertial Navigation System (INS) units. Vehicle command and control was managed by the Vehicle Management Computer (VMC) and supported by distributed Subsystem Management Units (SMUs) which were responsible for valve actuation, pyro initiation, and subsystem health monitoring. The main engines and OMS each had their own control units that performed the same functions. Data communications are conducted using a triple-redundant 1553 data bus.
The K-1 vehicle avionics included a Tracking Data Relay Satellite System (TDRSS) transceiver for communicating vehicle status to the ground and for updating OV wind data for re-entry targeting. TDRSS was also used to provide the customer with payload deployment data to support ground station acquisition. Kistler was responsible for scheduling the use of TDRSS with NASA.
All mission software was fully verified prior to flight in a hardware-in-the-loop system operated by Draper Laboratory.
Kistler settled on Woomera, Australia (31.08° South latitude, 136.66° East longitude) as its primary launch site. A spaceport in the continental United States, possibly at the Nevada Test Site near Las Vegas, Nevada, USA (37.17° North latitude, 116.27° West longitude), was a future possibility as of 2006. Either launch site allowed orbital inclinations of 45 to 60 degrees or 84 to 99 degrees to be reached.
LEO Payload: 4,500 kg (9,900 lb) to a 200 km orbit at 45.00 degrees. Payload: 2,000 kg (4,400 lb) to a 900 km 60 deg orbit. Launch Price $: 17.000 million in 1997 dollars.
Stage Data - Kistler K-1
Gross mass: 382,300 kg (842,800 lb).
Payload: 4,500 kg (9,900 lb).
Height: 36.90 m (121.00 ft).
Diameter: 6.70 m (21.90 ft).
Thrust: 4,540.00 kN (1,020,630 lbf).
Apogee: 200 km (120 mi).